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

GOE 342 AIRFOIL (goe342-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 342 AIRFOIL (goe342-il)
Reynolds number: 200,000
Max Cl/Cd: 99.45 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe342-il-200000.txt
Download as CSV file: xf-goe342-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 342 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.2994   0.11231   0.10879  -0.0319   1.0000   0.0240
  -8.500  -0.3075   0.11178   0.10834  -0.0308   1.0000   0.0241
  -7.500  -0.3358   0.10363   0.10047  -0.0181   1.0000   0.0245
  -7.250  -0.3290   0.10032   0.09717  -0.0181   0.9983   0.0248
  -7.000  -0.3056   0.09600   0.09283  -0.0231   0.9946   0.0253
  -6.750  -0.2807   0.09189   0.08871  -0.0290   0.9892   0.0260
  -6.500  -0.2526   0.08776   0.08456  -0.0359   0.9848   0.0270
  -6.250  -0.2244   0.08386   0.08063  -0.0430   0.9789   0.0282
  -6.000  -0.1547   0.08057   0.07720  -0.0675   0.9716   0.0297
  -5.750  -0.1339   0.07538   0.07201  -0.0706   0.9673   0.0301
  -5.500  -0.1221   0.07129   0.06794  -0.0692   0.9628   0.0306
  -5.250  -0.0949   0.06750   0.06411  -0.0732   0.9594   0.0315
  -5.000  -0.0577   0.06360   0.06017  -0.0805   0.9571   0.0329
  -4.750  -0.0222   0.06018   0.05668  -0.0875   0.9513   0.0350
  -4.500   0.0545   0.05580   0.05202  -0.1064   0.9468   0.0373
  -4.250   0.0731   0.05178   0.04804  -0.1070   0.9442   0.0382
  -4.000   0.1070   0.04850   0.04471  -0.1113   0.9420   0.0398
  -3.750   0.1397   0.04577   0.04190  -0.1154   0.9353   0.0420
  -3.500   0.2087   0.04219   0.03791  -0.1275   0.9314   0.0467
  -3.250   0.2349   0.03881   0.03457  -0.1295   0.9280   0.0481
  -3.000   0.2643   0.03660   0.03231  -0.1317   0.9223   0.0505
  -2.750   0.3204   0.03572   0.03088  -0.1375   0.9172   0.0572
  -2.500   0.3459   0.03149   0.02677  -0.1398   0.9135   0.0593
  -2.250   0.3735   0.02985   0.02510  -0.1410   0.9081   0.0637
  -2.000   0.4106   0.02792   0.02290  -0.1435   0.9022   0.0720
  -1.750   0.4410   0.02639   0.02132  -0.1447   0.8979   0.0778
  -1.500   0.4727   0.02489   0.01963  -0.1460   0.8917   0.0865
  -1.250   0.5050   0.02369   0.01820  -0.1470   0.8857   0.0986
  -1.000   0.5353   0.02252   0.01687  -0.1475   0.8790   0.1119
  -0.750   0.5628   0.02127   0.01554  -0.1476   0.8701   0.1265
  -0.500   0.5895   0.02014   0.01441  -0.1477   0.8625   0.1447
  -0.250   0.6166   0.01929   0.01354  -0.1480   0.8561   0.1764
   0.000   0.6442   0.01834   0.01257  -0.1483   0.8512   0.2188
   0.750   0.7456   0.01525   0.00822  -0.1470   0.8311   0.0806
   1.000   0.7747   0.01439   0.00716  -0.1464   0.8240   0.0722
   1.250   0.8020   0.01396   0.00669  -0.1459   0.8160   0.0715
   1.500   0.8298   0.01349   0.00620  -0.1454   0.8084   0.0719
   1.750   0.8566   0.01324   0.00597  -0.1448   0.7992   0.0704
   2.000   0.8843   0.01290   0.00564  -0.1443   0.7908   0.0699
   2.250   0.9112   0.01258   0.00537  -0.1438   0.7794   0.0705
   2.500   0.9378   0.01235   0.00514  -0.1431   0.7662   0.0726
   2.750   0.9641   0.01221   0.00500  -0.1423   0.7516   0.0772
   3.000   0.9905   0.01210   0.00491  -0.1415   0.7359   0.0919
   3.250   1.0118   0.01069   0.00494  -0.1399   0.7194   1.0000
   3.500   1.0371   0.01074   0.00486  -0.1388   0.6959   1.0000
   3.750   1.0616   0.01079   0.00480  -0.1377   0.6630   1.0000
   4.000   1.0850   0.01091   0.00474  -0.1363   0.6074   1.0000
   4.250   1.1050   0.01143   0.00475  -0.1345   0.5317   1.0000
   4.500   1.1257   0.01217   0.00510  -0.1330   0.4848   1.0000
   4.750   1.1478   0.01282   0.00551  -0.1320   0.4540   1.0000
   5.000   1.1707   0.01340   0.00594  -0.1312   0.4301   1.0000
   5.250   1.1942   0.01390   0.00635  -0.1304   0.4111   1.0000
   5.500   1.2180   0.01438   0.00677  -0.1297   0.3962   1.0000
   5.750   1.2417   0.01484   0.00719  -0.1290   0.3818   1.0000
   6.000   1.2654   0.01529   0.00765  -0.1283   0.3687   1.0000
   6.250   1.2888   0.01576   0.00811  -0.1276   0.3561   1.0000
   6.500   1.3119   0.01622   0.00857  -0.1268   0.3429   1.0000
   6.750   1.3347   0.01663   0.00903  -0.1260   0.3272   1.0000
   7.000   1.3570   0.01700   0.00951  -0.1250   0.3072   1.0000
   7.250   1.3779   0.01742   0.00994  -0.1239   0.2784   1.0000
   7.500   1.3959   0.01807   0.01039  -0.1224   0.2140   1.0000
   7.750   1.4060   0.01971   0.01143  -0.1200   0.1452   1.0000
   8.000   1.4195   0.02103   0.01261  -0.1179   0.1235   1.0000
   8.250   1.4339   0.02221   0.01379  -0.1160   0.1100   1.0000
   8.500   1.4475   0.02341   0.01496  -0.1139   0.0988   1.0000
   8.750   1.4635   0.02433   0.01592  -0.1122   0.0871   1.0000
   9.000   1.4785   0.02527   0.01688  -0.1104   0.0737   1.0000
   9.250   1.4897   0.02654   0.01815  -0.1079   0.0589   1.0000
   9.500   1.4922   0.02829   0.01990  -0.1042   0.0492   1.0000
   9.750   1.5009   0.02957   0.02127  -0.1014   0.0428   1.0000
  10.000   1.5030   0.03148   0.02321  -0.0980   0.0390   1.0000
  10.250   1.5114   0.03292   0.02479  -0.0954   0.0360   1.0000
  10.500   1.5179   0.03454   0.02648  -0.0929   0.0338   1.0000
  10.750   1.5210   0.03669   0.02866  -0.0901   0.0320   1.0000
  11.000   1.5270   0.03899   0.03108  -0.0876   0.0307   1.0000
  11.250   1.5355   0.04076   0.03305  -0.0854   0.0293   1.0000
  11.500   1.5428   0.04260   0.03505  -0.0834   0.0278   1.0000
  11.750   1.5487   0.04452   0.03708  -0.0814   0.0265   1.0000
  12.000   1.5546   0.04671   0.03938  -0.0795   0.0257   1.0000
  12.250   1.5606   0.04936   0.04212  -0.0776   0.0249   1.0000
  12.500   1.5674   0.05333   0.04632  -0.0756   0.0242   1.0000
  12.750   1.5665   0.05647   0.04974  -0.0734   0.0239   1.0000
  13.000   1.5627   0.05978   0.05336  -0.0714   0.0237   1.0000
  13.250   1.5557   0.06347   0.05734  -0.0696   0.0236   1.0000
  13.500   1.5456   0.06748   0.06165  -0.0681   0.0234   1.0000
  13.750   1.5328   0.07185   0.06630  -0.0671   0.0233   1.0000
  14.000   1.5178   0.07660   0.07132  -0.0666   0.0232   1.0000
  14.250   1.5007   0.08181   0.07680  -0.0666   0.0231   1.0000
  14.500   1.4822   0.08745   0.08269  -0.0674   0.0231   1.0000
  14.750   1.4623   0.09361   0.08909  -0.0690   0.0231   1.0000
  15.000   1.4415   0.10040   0.09611  -0.0715   0.0231   1.0000
  15.250   1.4201   0.10795   0.10388  -0.0750   0.0232   1.0000
  15.500   1.3986   0.11614   0.11228  -0.0795   0.0234   1.0000
  15.750   1.3770   0.12503   0.12136  -0.0849   0.0236   1.0000
  16.000   1.3551   0.13462   0.13109  -0.0909   0.0239   1.0000
  16.250   1.0817   0.13434   0.13127  -0.0726   0.0270   1.0000
  16.500   1.0013   0.15506   0.15217  -0.0857   0.0298   1.0000
<< Back to GOE 342 AIRFOIL (goe342-il)

Polar data table (+)

Polar graphs


<< Back to GOE 342 AIRFOIL (goe342-il)