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

GOE 652 AIRFOIL (goe652-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 652 AIRFOIL (goe652-il)
Reynolds number: 100,000
Max Cl/Cd: 53.13 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe652-il-100000-n5.txt
Download as CSV file: xf-goe652-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 652 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250   0.2396   0.10650   0.10022  -0.1278   0.8810   0.1054
  -9.000   0.2537   0.10395   0.09765  -0.1301   0.8751   0.1055
  -8.750   0.2638   0.09866   0.09225  -0.1338   0.8706   0.0894
  -8.500   0.2891   0.09569   0.08925  -0.1368   0.8673   0.0874
  -8.250   0.2927   0.09388   0.08745  -0.1362   0.8587   0.0860
  -8.000   0.3050   0.09158   0.08513  -0.1378   0.8526   0.0863
  -7.750   0.3211   0.08897   0.08248  -0.1404   0.8478   0.0870
  -7.500   0.3318   0.08675   0.08025  -0.1416   0.8419   0.0869
  -7.250   0.3355   0.08498   0.07849  -0.1412   0.8339   0.0863
  -7.000   0.3456   0.08258   0.07606  -0.1425   0.8281   0.0858
  -6.750   0.3591   0.07995   0.07338  -0.1448   0.8237   0.0856
  -6.500   0.3577   0.07852   0.07199  -0.1433   0.8158   0.0855
  -6.250   0.3616   0.07673   0.07020  -0.1430   0.8088   0.0858
  -6.000   0.3714   0.07457   0.06801  -0.1442   0.8034   0.0873
  -5.750   0.3718   0.07257   0.06600  -0.1441   0.7972   0.0882
  -5.500   0.3615   0.07111   0.06460  -0.1416   0.7883   0.0888
  -5.250   0.3565   0.06843   0.06189  -0.1416   0.7821   0.0894
  -5.000   0.3609   0.06598   0.05941  -0.1424   0.7777   0.0901
  -4.750   0.3450   0.06597   0.05951  -0.1370   0.7680   0.0904
  -4.500   0.3540   0.06484   0.05839  -0.1364   0.7623   0.0913
  -4.250   0.3706   0.06321   0.05673  -0.1376   0.7581   0.0929
  -4.000   0.3647   0.06198   0.05554  -0.1357   0.7507   0.0945
  -3.750   0.3496   0.05923   0.05283  -0.1354   0.7420   0.0966
  -3.500   0.3642   0.05609   0.04963  -0.1392   0.7375   0.0987
  -3.250   0.3934   0.05434   0.04784  -0.1424   0.7345   0.1005
  -3.000   0.3760   0.05369   0.04729  -0.1384   0.7240   0.1018
  -2.750   0.4327   0.04274   0.03606  -0.1646   0.7188   0.1150
  -2.250   0.5796   0.03487   0.02776  -0.1943   0.7124   0.1348
  -1.750   0.7180   0.02888   0.02124  -0.2219   0.6982   0.1597
  -1.250   0.8004   0.02761   0.01973  -0.2283   0.6875   0.1732
  -1.000   0.8145   0.02813   0.02037  -0.2250   0.6791   0.1750
  -0.500   0.8984   0.02676   0.01868  -0.2314   0.6672   0.1893
  -0.250   0.9153   0.02717   0.01921  -0.2291   0.6594   0.1920
   0.000   0.9559   0.02662   0.01849  -0.2321   0.6540   0.1999
   0.250   0.9972   0.02608   0.01779  -0.2347   0.6495   0.2065
   0.500   1.0153   0.02644   0.01826  -0.2328   0.6418   0.2098
   0.750   1.0497   0.02614   0.01783  -0.2345   0.6349   0.2172
   1.000   1.0889   0.02568   0.01720  -0.2365   0.6295   0.2240
   1.250   1.1089   0.02598   0.01759  -0.2350   0.6218   0.2274
   1.500   1.1358   0.02601   0.01760  -0.2347   0.6143   0.2331
   1.750   1.1780   0.02548   0.01679  -0.2376   0.6085   0.2420
   2.000   1.1966   0.02583   0.01726  -0.2358   0.6007   0.2451
   2.250   1.2205   0.02600   0.01745  -0.2348   0.5931   0.2497
   2.500   1.2569   0.02574   0.01702  -0.2361   0.5871   0.2572
   2.750   1.2794   0.02595   0.01720  -0.2355   0.5782   0.2632
   3.000   1.3036   0.02610   0.01739  -0.2345   0.5704   0.2670
   3.250   1.3302   0.02620   0.01746  -0.2340   0.5627   0.2721
   3.500   1.3518   0.02645   0.01768  -0.2330   0.5531   0.2782
   3.750   1.3853   0.02633   0.01738  -0.2338   0.5453   0.2852
   4.000   1.3994   0.02683   0.01799  -0.2314   0.5350   0.2887
   4.250   1.4250   0.02697   0.01808  -0.2307   0.5264   0.2940
   4.500   1.4439   0.02736   0.01846  -0.2293   0.5168   0.3001
   4.750   1.4668   0.02761   0.01860  -0.2284   0.5075   0.3069
   5.000   1.4844   0.02801   0.01904  -0.2265   0.4982   0.3109
   5.250   1.5008   0.02848   0.01951  -0.2246   0.4881   0.3158
   5.500   1.5202   0.02892   0.01989  -0.2232   0.4785   0.3226
   5.750   1.5366   0.02948   0.02040  -0.2214   0.4680   0.3292
   6.000   1.5529   0.03005   0.02097  -0.2195   0.4584   0.3340
   6.250   1.5667   0.03075   0.02167  -0.2174   0.4480   0.3396
   6.500   1.5830   0.03142   0.02225  -0.2158   0.4386   0.3467
   6.750   1.5971   0.03220   0.02303  -0.2139   0.4294   0.3527
   7.000   1.6143   0.03287   0.02365  -0.2123   0.4219   0.3592
   7.250   1.6274   0.03381   0.02462  -0.2105   0.4140   0.3665
   7.500   1.6451   0.03453   0.02528  -0.2091   0.4074   0.3735
   7.750   1.6602   0.03541   0.02617  -0.2075   0.4009   0.3800
   8.000   1.6727   0.03646   0.02722  -0.2057   0.3937   0.3881
   8.250   1.6894   0.03727   0.02798  -0.2042   0.3874   0.3958
   8.500   1.7037   0.03827   0.02900  -0.2027   0.3816   0.4035
   8.750   1.7158   0.03946   0.03020  -0.2010   0.3756   0.4120
   9.000   1.7312   0.04040   0.03115  -0.1995   0.3704   0.4195
   9.250   1.7529   0.04103   0.03167  -0.1986   0.3658   0.4303
   9.500   1.7618   0.04247   0.03326  -0.1968   0.3609   0.4378
   9.750   1.7734   0.04378   0.03465  -0.1952   0.3560   0.4468
  10.000   1.7885   0.04486   0.03575  -0.1939   0.3515   0.4567
  10.250   1.8096   0.04555   0.03639  -0.1930   0.3473   0.4687
  10.500   1.8210   0.04693   0.03787  -0.1915   0.3429   0.4786
  10.750   1.8272   0.04871   0.03978  -0.1897   0.3379   0.4887
  11.000   1.8368   0.05024   0.04140  -0.1882   0.3330   0.4997
  11.250   1.8545   0.05117   0.04231  -0.1872   0.3283   0.5145
  11.500   1.8685   0.05243   0.04362  -0.1860   0.3238   0.5293
  11.750   1.8682   0.05485   0.04627  -0.1841   0.3188   0.5414
  12.000   1.8741   0.05681   0.04836  -0.1826   0.3139   0.5581
  12.250   1.8865   0.05819   0.04982  -0.1814   0.3093   0.5821
  12.500   1.9029   0.05924   0.05091  -0.1804   0.3045   0.6207
  12.750   1.8919   0.06283   0.05484  -0.1785   0.2985   0.6608
  13.000   1.8892   0.06546   0.05770  -0.1770   0.2924   1.0000
  13.250   1.9034   0.06669   0.05879  -0.1760   0.2869   1.0000
  13.500   1.8883   0.07121   0.06355  -0.1746   0.2803   1.0000
  13.750   1.8817   0.07490   0.06737  -0.1736   0.2736   1.0000
  14.000   1.8921   0.07652   0.06890  -0.1727   0.2681   1.0000
  14.250   1.8727   0.08218   0.07482  -0.1721   0.2611   1.0000
  14.500   1.8672   0.08610   0.07884  -0.1716   0.2549   1.0000
  14.750   1.8750   0.08819   0.08088  -0.1710   0.2500   1.0000
  15.000   1.8538   0.09470   0.08766  -0.1713   0.2429   1.0000
  15.250   1.8518   0.09835   0.09135  -0.1713   0.2372   1.0000
  15.500   1.8486   0.10226   0.09533  -0.1714   0.2319   1.0000
  15.750   1.8321   0.10844   0.10170  -0.1724   0.2254   1.0000
  16.000   1.8370   0.11110   0.10434  -0.1724   0.2207   1.0000
<< Back to GOE 652 AIRFOIL (goe652-il)

Polar data table (+)

Polar graphs


<< Back to GOE 652 AIRFOIL (goe652-il)