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GOE 499 AIRFOIL (goe499-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 499 AIRFOIL (goe499-il)
Reynolds number: 200,000
Max Cl/Cd: 102.16 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe499-il-200000.txt
Download as CSV file: xf-goe499-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 499 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3149   0.10306   0.09980  -0.0350   0.9896   0.0379
  -7.500  -0.3034   0.10027   0.09702  -0.0402   0.9854   0.0400
  -7.250  -0.3027   0.09909   0.09586  -0.0453   0.9783   0.0409
  -7.000  -0.2824   0.09548   0.09225  -0.0539   0.9727   0.0413
  -6.750  -0.2702   0.08894   0.08573  -0.0568   0.9707   0.0423
  -6.500  -0.2589   0.08654   0.08331  -0.0528   0.9690   0.0445
  -6.250  -0.2522   0.08427   0.08104  -0.0531   0.9634   0.0465
  -6.000  -0.2319   0.08072   0.07747  -0.0584   0.9593   0.0497
  -5.750  -0.1931   0.07612   0.07281  -0.0784   0.9517   0.0540
  -5.500  -0.1853   0.07106   0.06776  -0.0789   0.9478   0.0553
  -5.250  -0.1713   0.06895   0.06563  -0.0766   0.9455   0.0571
  -5.000  -0.1414   0.06593   0.06256  -0.0816   0.9435   0.0630
  -4.750  -0.1078   0.06010   0.05664  -0.0948   0.9372   0.0688
  -4.500  -0.0458   0.05524   0.05150  -0.1117   0.9338   0.0801
  -4.250  -0.0287   0.05144   0.04776  -0.1115   0.9314   0.0820
  -4.000  -0.0010   0.04975   0.04605  -0.1121   0.9294   0.0860
  -3.750   0.0429   0.04509   0.04119  -0.1218   0.9238   0.0970
  -3.250   0.1993   0.02592   0.01991  -0.1537   0.9222   0.0677
  -3.000   0.2421   0.02330   0.01694  -0.1573   0.9195   0.0654
  -2.750   0.2875   0.02108   0.01427  -0.1610   0.9175   0.0632
  -2.500   0.3331   0.01956   0.01241  -0.1643   0.9160   0.0628
  -2.250   0.3651   0.01874   0.01143  -0.1650   0.9114   0.0634
  -2.000   0.3952   0.01813   0.01073  -0.1654   0.9068   0.0644
  -1.750   0.4327   0.01759   0.01010  -0.1672   0.9045   0.0669
  -1.500   0.4733   0.01683   0.00938  -0.1697   0.9029   0.0707
  -1.250   0.5144   0.01631   0.00885  -0.1722   0.9015   0.0752
  -1.000   0.5358   0.01626   0.00881  -0.1709   0.8950   0.0818
  -0.750   0.5708   0.01571   0.00855  -0.1723   0.8918   0.1493
  -0.500   0.6081   0.01535   0.00839  -0.1740   0.8894   0.2393
  -0.250   0.6466   0.01499   0.00814  -0.1759   0.8874   0.2866
   0.000   0.6689   0.01498   0.00828  -0.1749   0.8804   0.3289
   0.250   0.7024   0.01455   0.00818  -0.1759   0.8766   0.4158
   0.500   0.7376   0.01401   0.00803  -0.1770   0.8740   0.5649
   0.750   0.7615   0.01321   0.00796  -0.1757   0.8685   1.0000
   1.000   0.7919   0.01312   0.00779  -0.1759   0.8629   1.0000
   1.250   0.8280   0.01285   0.00745  -0.1770   0.8596   1.0000
   1.500   0.8528   0.01290   0.00748  -0.1762   0.8515   1.0000
   1.750   0.8868   0.01260   0.00717  -0.1768   0.8464   1.0000
   2.000   0.9137   0.01252   0.00708  -0.1762   0.8378   1.0000
   2.250   0.9472   0.01218   0.00674  -0.1766   0.8314   1.0000
   2.500   0.9727   0.01213   0.00672  -0.1757   0.8212   1.0000
   2.750   1.0020   0.01195   0.00659  -0.1755   0.8124   1.0000
   3.000   1.0292   0.01159   0.00622  -0.1744   0.7961   1.0000
   3.250   1.0570   0.01112   0.00571  -0.1733   0.7742   1.0000
   3.500   1.0789   0.01085   0.00539  -0.1711   0.7375   1.0000
   3.750   1.1023   0.01079   0.00526  -0.1695   0.6981   1.0000
   4.000   1.1225   0.01102   0.00509  -0.1671   0.6197   1.0000
   4.250   1.1323   0.01212   0.00543  -0.1631   0.5113   1.0000
   4.500   1.1381   0.01358   0.00615  -0.1589   0.3953   1.0000
   4.750   1.1276   0.01689   0.00773  -0.1530   0.1076   1.0000
   5.000   1.1404   0.01836   0.00883  -0.1504   0.0435   1.0000
   5.250   1.1580   0.01933   0.00992  -0.1484   0.0376   1.0000
   5.500   1.1765   0.02014   0.01085  -0.1466   0.0343   1.0000
   5.750   1.1924   0.02113   0.01195  -0.1444   0.0323   1.0000
   6.000   1.2060   0.02228   0.01318  -0.1419   0.0310   1.0000
   6.250   1.2176   0.02355   0.01452  -0.1390   0.0301   1.0000
   6.500   1.2290   0.02497   0.01598  -0.1361   0.0296   1.0000
   6.750   1.2436   0.02653   0.01760  -0.1338   0.0293   1.0000
   7.000   1.2676   0.02900   0.02004  -0.1333   0.0278   1.0000
   7.250   1.3100   0.03248   0.02355  -0.1355   0.0272   1.0000
   7.500   1.3386   0.03401   0.02519  -0.1351   0.0277   1.0000
   7.750   1.3675   0.03538   0.02689  -0.1340   0.0302   1.0000
  12.500   1.2673   0.11986   0.11684  -0.0987   0.0458   1.0000
  12.750   1.2451   0.12819   0.12527  -0.1036   0.0464   1.0000
  13.000   0.9902   0.15711   0.15471  -0.1186   0.0681   1.0000
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