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=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 342 AIRFOIL (goe342-il)
Reynolds number: 50,000
Max Cl/Cd: 47.81 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe342-il-50000-n5.txt
Download as CSV file: xf-goe342-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.3064   0.10627   0.09990  -0.0232   1.0000   0.0580
  -7.000  -0.3113   0.10474   0.09847  -0.0216   1.0000   0.0593
  -6.750  -0.3163   0.10333   0.09715  -0.0202   1.0000   0.0608
  -6.500  -0.3207   0.10208   0.09598  -0.0195   1.0000   0.0627
  -6.250  -0.3234   0.10143   0.09539  -0.0208   1.0000   0.0646
  -6.000  -0.3191   0.10149   0.09550  -0.0260   1.0000   0.0655
  -5.750  -0.3149   0.09602   0.09008  -0.0220   0.9981   0.0677
  -5.500  -0.2950   0.09228   0.08632  -0.0252   0.9934   0.0722
  -5.250  -0.2456   0.09042   0.08429  -0.0426   0.9860   0.0775
  -5.000  -0.2394   0.08536   0.07931  -0.0396   0.9823   0.0794
  -4.750  -0.2191   0.08169   0.07559  -0.0419   0.9781   0.0841
  -4.500  -0.1775   0.07869   0.07246  -0.0534   0.9721   0.0913
  -4.250  -0.1614   0.07492   0.06869  -0.0534   0.9684   0.0978
  -4.000  -0.1218   0.07175   0.06540  -0.0627   0.9633   0.1051
  -3.750  -0.0993   0.06839   0.06200  -0.0647   0.9593   0.1117
  -3.500  -0.0598   0.06502   0.05851  -0.0723   0.9555   0.1202
  -3.250  -0.0240   0.06223   0.05556  -0.0784   0.9508   0.1333
  -3.000   0.0092   0.05921   0.05246  -0.0829   0.9473   0.1488
  -2.750   0.0436   0.05640   0.04955  -0.0875   0.9438   0.1644
  -2.500   0.0746   0.05385   0.04691  -0.0911   0.9394   0.1803
  -2.250   0.1091   0.05125   0.04421  -0.0949   0.9359   0.1991
  -2.000   0.1469   0.04877   0.04163  -0.0994   0.9332   0.2293
  -1.750   0.1701   0.04673   0.03960  -0.1005   0.9281   0.2726
  -1.250   0.3152   0.04136   0.03268  -0.1208   0.9223   0.1021
  -1.000   0.3632   0.03912   0.03005  -0.1254   0.9198   0.0884
  -0.750   0.3953   0.03771   0.02845  -0.1271   0.9135   0.0853
  -0.500   0.4402   0.03623   0.02662  -0.1308   0.9081   0.0855
  -0.250   0.4810   0.03489   0.02496  -0.1334   0.9004   0.0837
   0.000   0.5309   0.03331   0.02294  -0.1370   0.8921   0.0803
   0.250   0.5723   0.03214   0.02134  -0.1388   0.8805   0.0783
   0.500   0.6162   0.03095   0.01983  -0.1410   0.8702   0.0775
   0.750   0.6573   0.02990   0.01852  -0.1426   0.8614   0.0775
   1.000   0.6866   0.02934   0.01784  -0.1425   0.8517   0.0790
   1.250   0.7246   0.02861   0.01702  -0.1438   0.8453   0.0841
   1.500   0.7507   0.02835   0.01668  -0.1431   0.8347   0.0873
   1.750   0.7849   0.02780   0.01605  -0.1435   0.8269   0.0894
   2.000   0.8151   0.02741   0.01567  -0.1434   0.8166   0.0924
   2.250   0.8438   0.02715   0.01541  -0.1431   0.8054   0.0972
   2.500   0.8768   0.02673   0.01500  -0.1433   0.7955   0.1052
   2.750   0.9087   0.02629   0.01473  -0.1433   0.7843   0.1264
   3.000   0.9311   0.02488   0.01474  -0.1418   0.7710   1.0000
   3.250   0.9582   0.02487   0.01457  -0.1408   0.7569   1.0000
   3.500   0.9853   0.02483   0.01447  -0.1398   0.7419   1.0000
   3.750   1.0122   0.02477   0.01437  -0.1388   0.7257   1.0000
   4.000   1.0390   0.02469   0.01428  -0.1377   0.7082   1.0000
   4.250   1.0628   0.02476   0.01436  -0.1363   0.6875   1.0000
   4.500   1.0855   0.02488   0.01457  -0.1348   0.6639   1.0000
   4.750   1.1081   0.02502   0.01476  -0.1333   0.6381   1.0000
   5.000   1.1327   0.02509   0.01485  -0.1320   0.6117   1.0000
   5.250   1.1586   0.02513   0.01492  -0.1308   0.5859   1.0000
   5.500   1.1846   0.02525   0.01498  -0.1297   0.5605   1.0000
   5.750   1.2105   0.02544   0.01505  -0.1286   0.5357   1.0000
   6.000   1.2349   0.02583   0.01533  -0.1273   0.5113   1.0000
   6.250   1.2575   0.02641   0.01587  -0.1261   0.4877   1.0000
   6.500   1.2795   0.02708   0.01650  -0.1248   0.4661   1.0000
   6.750   1.3004   0.02782   0.01721  -0.1234   0.4444   1.0000
   7.000   1.3194   0.02863   0.01805  -0.1219   0.4221   1.0000
   7.500   1.3541   0.03035   0.01990  -0.1184   0.3775   1.0000
   7.750   1.3697   0.03128   0.02085  -0.1164   0.3550   1.0000
   8.000   1.3833   0.03226   0.02198  -0.1143   0.3319   1.0000
   8.250   1.3954   0.03331   0.02312  -0.1119   0.3081   1.0000
   8.500   1.4053   0.03444   0.02449  -0.1094   0.2815   1.0000
   8.750   1.4124   0.03568   0.02587  -0.1066   0.2509   1.0000
   9.000   1.4165   0.03707   0.02728  -0.1034   0.2193   1.0000
   9.250   1.4198   0.03869   0.02881  -0.1004   0.1923   1.0000
   9.500   1.4235   0.04057   0.03056  -0.0976   0.1736   1.0000
   9.750   1.4278   0.04262   0.03252  -0.0951   0.1580   1.0000
  10.000   1.4318   0.04483   0.03468  -0.0927   0.1445   1.0000
  10.250   1.4348   0.04721   0.03709  -0.0904   0.1312   1.0000
  10.500   1.4349   0.04978   0.03971  -0.0883   0.1172   1.0000
  10.750   1.4333   0.05249   0.04255  -0.0863   0.1042   1.0000
  11.000   1.4325   0.05531   0.04555  -0.0844   0.0936   1.0000
  11.250   1.4301   0.05830   0.04860  -0.0827   0.0852   1.0000
  11.500   1.4276   0.06142   0.05197  -0.0812   0.0765   1.0000
  11.750   1.4253   0.06472   0.05543  -0.0797   0.0697   1.0000
  12.000   1.4221   0.06811   0.05897  -0.0785   0.0639   1.0000
  12.250   1.4187   0.07165   0.06261  -0.0774   0.0593   1.0000
  12.500   1.4156   0.07543   0.06670  -0.0765   0.0547   1.0000
  12.750   1.4114   0.07919   0.07064  -0.0759   0.0515   1.0000
  13.000   1.4104   0.08266   0.07412  -0.0751   0.0490   1.0000
  13.250   1.4053   0.08725   0.07909  -0.0747   0.0471   1.0000
  13.500   1.3964   0.09236   0.08456  -0.0751   0.0454   1.0000
  13.750   1.3852   0.09782   0.09032  -0.0761   0.0441   1.0000
  14.000   1.3726   0.10367   0.09644  -0.0778   0.0431   1.0000
  14.250   1.3587   0.11007   0.10307  -0.0802   0.0424   1.0000
  14.500   1.3433   0.11720   0.11044  -0.0836   0.0419   1.0000
  14.750   1.3251   0.12561   0.11910  -0.0882   0.0419   1.0000
  15.000   1.3015   0.13632   0.13006  -0.0947   0.0428   1.0000
  15.250   1.2740   0.14944   0.14334  -0.1033   0.0443   1.0000
<< Back to GOE 342 AIRFOIL (goe342-il)

Polar data table (+)

Polar graphs


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