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

GOE 601 AIRFOIL (goe601-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 601 AIRFOIL (goe601-il)
Reynolds number: 200,000
Max Cl/Cd: 57.44 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe601-il-200000-n5.txt
Download as CSV file: xf-goe601-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 601 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -17.250  -0.7216   0.10208   0.09731  -0.0686   1.0000   0.0314
 -17.000  -0.7645   0.09072   0.08568  -0.0753   1.0000   0.0314
 -16.750  -0.7921   0.08280   0.07755  -0.0799   1.0000   0.0315
 -16.500  -0.8163   0.07599   0.07052  -0.0837   1.0000   0.0317
 -16.250  -0.8341   0.07061   0.06496  -0.0863   1.0000   0.0318
 -16.000  -0.8377   0.06751   0.06184  -0.0871   1.0000   0.0322
 -15.750  -0.8428   0.06434   0.05862  -0.0879   1.0000   0.0325
 -15.500  -0.8482   0.06122   0.05544  -0.0887   1.0000   0.0328
 -15.250  -0.8542   0.05814   0.05226  -0.0893   1.0000   0.0331
 -15.000  -0.8598   0.05526   0.04931  -0.0897   1.0000   0.0335
 -14.750  -0.8654   0.05248   0.04644  -0.0899   1.0000   0.0338
 -14.500  -0.8708   0.04989   0.04375  -0.0898   1.0000   0.0343
 -14.250  -0.8762   0.04744   0.04119  -0.0893   1.0000   0.0347
 -14.000  -0.8816   0.04519   0.03882  -0.0885   1.0000   0.0351
 -13.750  -0.8870   0.04311   0.03663  -0.0874   1.0000   0.0356
 -13.500  -0.8921   0.04124   0.03464  -0.0858   1.0000   0.0360
 -13.250  -0.8978   0.03953   0.03280  -0.0838   1.0000   0.0364
 -13.000  -0.9036   0.03801   0.03115  -0.0815   1.0000   0.0368
 -12.750  -0.9085   0.03664   0.02973  -0.0787   1.0000   0.0372
 -12.500  -0.9068   0.03532   0.02839  -0.0769   0.9993   0.0376
 -12.250  -0.8865   0.03377   0.02680  -0.0785   0.9955   0.0385
 -12.000  -0.8667   0.03236   0.02532  -0.0798   0.9911   0.0393
 -11.750  -0.8462   0.03105   0.02392  -0.0810   0.9862   0.0403
 -11.500  -0.8225   0.02982   0.02258  -0.0824   0.9825   0.0412
 -11.250  -0.8042   0.02872   0.02136  -0.0825   0.9762   0.0421
 -11.000  -0.7805   0.02765   0.02016  -0.0835   0.9719   0.0429
 -10.750  -0.7638   0.02651   0.01902  -0.0832   0.9652   0.0439
 -10.250  -0.7188   0.02463   0.01704  -0.0839   0.9549   0.0463
 -10.000  -0.7008   0.02384   0.01616  -0.0830   0.9471   0.0475
  -9.750  -0.6727   0.02296   0.01518  -0.0841   0.9434   0.0487
  -9.500  -0.6560   0.02220   0.01438  -0.0828   0.9346   0.0500
  -9.250  -0.6319   0.02139   0.01354  -0.0831   0.9292   0.0519
  -9.000  -0.6124   0.02076   0.01286  -0.0821   0.9216   0.0539
  -8.750  -0.5900   0.02014   0.01216  -0.0816   0.9148   0.0564
  -8.500  -0.5676   0.01949   0.01152  -0.0812   0.9085   0.0591
  -8.250  -0.5477   0.01897   0.01096  -0.0801   0.9005   0.0630
  -8.000  -0.5203   0.01838   0.01040  -0.0805   0.8963   0.0696
  -7.750  -0.5004   0.01799   0.01004  -0.0793   0.8882   0.0780
  -7.500  -0.4726   0.01759   0.00961  -0.0796   0.8832   0.0890
  -7.250  -0.4444   0.01723   0.00919  -0.0799   0.8786   0.0983
  -7.000  -0.4208   0.01695   0.00884  -0.0793   0.8720   0.1056
  -6.750  -0.3936   0.01653   0.00837  -0.0794   0.8673   0.1119
  -6.500  -0.3670   0.01617   0.00795  -0.0794   0.8624   0.1182
  -6.250  -0.3438   0.01589   0.00761  -0.0787   0.8563   0.1237
  -6.000  -0.3178   0.01548   0.00719  -0.0786   0.8516   0.1307
  -5.750  -0.2907   0.01515   0.00681  -0.0786   0.8472   0.1383
  -5.500  -0.2701   0.01482   0.00653  -0.0774   0.8410   0.1490
  -5.250  -0.2461   0.01443   0.00620  -0.0769   0.8362   0.1665
  -5.000  -0.2206   0.01395   0.00588  -0.0768   0.8322   0.2049
  -4.750  -0.2013   0.01366   0.00574  -0.0753   0.8262   0.2461
  -4.500  -0.1781   0.01341   0.00557  -0.0746   0.8212   0.2803
  -4.250  -0.1523   0.01315   0.00537  -0.0743   0.8171   0.3085
  -4.000  -0.1280   0.01297   0.00523  -0.0736   0.8125   0.3300
  -3.750  -0.1057   0.01284   0.00514  -0.0726   0.8072   0.3496
  -3.500  -0.0814   0.01268   0.00504  -0.0719   0.8026   0.3728
  -3.250  -0.0545   0.01254   0.00496  -0.0717   0.7988   0.4013
  -3.000  -0.0317   0.01249   0.00496  -0.0707   0.7939   0.4237
  -2.750  -0.0081   0.01246   0.00495  -0.0698   0.7888   0.4449
  -2.500   0.0182   0.01241   0.00489  -0.0694   0.7841   0.4643
  -2.250   0.0457   0.01236   0.00481  -0.0693   0.7791   0.4818
  -2.000   0.0674   0.01233   0.00480  -0.0679   0.7718   0.4955
  -1.750   0.0940   0.01228   0.00471  -0.0676   0.7658   0.5079
  -1.500   0.1201   0.01223   0.00465  -0.0671   0.7603   0.5200
  -1.250   0.1427   0.01218   0.00464  -0.0660   0.7532   0.5312
  -1.000   0.1694   0.01212   0.00457  -0.0656   0.7477   0.5427
  -0.750   0.1944   0.01207   0.00455  -0.0649   0.7422   0.5550
  -0.500   0.2169   0.01201   0.00455  -0.0638   0.7350   0.5687
  -0.250   0.2434   0.01193   0.00449  -0.0633   0.7292   0.5835
   0.000   0.2659   0.01189   0.00452  -0.0622   0.7222   0.5980
   0.250   0.2898   0.01183   0.00450  -0.0612   0.7150   0.6133
   0.750   0.3365   0.01171   0.00452  -0.0591   0.6998   0.6504
   1.000   0.3610   0.01163   0.00449  -0.0582   0.6917   0.6712
   1.250   0.3828   0.01157   0.00451  -0.0568   0.6819   0.6933
   1.500   0.4065   0.01151   0.00453  -0.0558   0.6735   0.7171
   1.750   0.4295   0.01146   0.00458  -0.0546   0.6645   0.7437
   2.000   0.4534   0.01141   0.00464  -0.0535   0.6552   0.7734
   2.250   0.4802   0.01138   0.00470  -0.0530   0.6446   0.8050
   2.500   0.5091   0.01142   0.00483  -0.0530   0.6330   0.8377
   2.750   0.5421   0.01149   0.00495  -0.0539   0.6200   0.8670
   3.000   0.5783   0.01161   0.00506  -0.0555   0.6039   0.8891
   3.250   0.6144   0.01176   0.00518  -0.0571   0.5854   0.9078
   3.500   0.6497   0.01195   0.00531  -0.0587   0.5644   0.9247
   3.750   0.6825   0.01218   0.00545  -0.0597   0.5389   0.9401
   4.000   0.7127   0.01250   0.00563  -0.0603   0.5073   0.9541
   4.250   0.7416   0.01291   0.00586  -0.0608   0.4684   0.9663
   4.500   0.7676   0.01343   0.00616  -0.0609   0.4284   0.9790
   4.750   0.7936   0.01402   0.00653  -0.0611   0.3909   0.9909
   5.000   0.8189   0.01460   0.00693  -0.0613   0.3633   1.0000
   5.250   0.8191   0.01490   0.00713  -0.0561   0.3494   1.0000
   5.500   0.8207   0.01521   0.00735  -0.0512   0.3378   1.0000
   5.750   0.8259   0.01554   0.00760  -0.0470   0.3273   1.0000
   6.000   0.8345   0.01586   0.00788  -0.0435   0.3178   1.0000
   6.250   0.8440   0.01623   0.00819  -0.0403   0.3095   1.0000
   6.500   0.8561   0.01657   0.00851  -0.0376   0.3018   1.0000
   6.750   0.8686   0.01693   0.00885  -0.0349   0.2942   1.0000
   7.000   0.8812   0.01733   0.00921  -0.0324   0.2867   1.0000
   7.250   0.8959   0.01767   0.00957  -0.0303   0.2795   1.0000
   7.500   0.9092   0.01811   0.00998  -0.0280   0.2733   1.0000
   7.750   0.9254   0.01846   0.01036  -0.0262   0.2677   1.0000
   8.000   0.9410   0.01884   0.01077  -0.0243   0.2612   1.0000
   8.250   0.9542   0.01933   0.01123  -0.0222   0.2539   1.0000
   8.500   0.9704   0.01972   0.01168  -0.0205   0.2463   1.0000
   8.750   0.9838   0.02023   0.01218  -0.0186   0.2393   1.0000
   9.000   0.9997   0.02067   0.01267  -0.0170   0.2310   1.0000
   9.250   1.0122   0.02126   0.01324  -0.0150   0.2224   1.0000
   9.500   1.0277   0.02176   0.01379  -0.0135   0.2129   1.0000
   9.750   1.0394   0.02243   0.01444  -0.0115   0.1984   1.0000
  10.000   1.0504   0.02319   0.01516  -0.0096   0.1821   1.0000
  10.250   1.0586   0.02415   0.01602  -0.0076   0.1599   1.0000
  10.500   1.0626   0.02545   0.01712  -0.0052   0.1259   1.0000
  10.750   1.0535   0.02772   0.01897  -0.0018   0.0740   1.0000
  11.000   1.0559   0.02933   0.02048   0.0004   0.0573   1.0000
  11.250   1.0620   0.03072   0.02185   0.0021   0.0503   1.0000
  11.500   1.0704   0.03197   0.02313   0.0036   0.0472   1.0000
  11.750   1.0771   0.03337   0.02457   0.0051   0.0447   1.0000
  12.000   1.0838   0.03481   0.02607   0.0065   0.0432   1.0000
  12.250   1.0908   0.03625   0.02760   0.0078   0.0419   1.0000
  12.500   1.0969   0.03781   0.02925   0.0090   0.0410   1.0000
  12.750   1.1017   0.03952   0.03104   0.0102   0.0398   1.0000
  13.000   1.1046   0.04145   0.03305   0.0113   0.0389   1.0000
  13.250   1.1061   0.04357   0.03524   0.0123   0.0383   1.0000
  13.500   1.1046   0.04605   0.03779   0.0132   0.0375   1.0000
  13.750   1.1071   0.04824   0.04009   0.0138   0.0370   1.0000
  14.000   1.1087   0.05060   0.04255   0.0143   0.0364   1.0000
  14.250   1.1091   0.05315   0.04520   0.0147   0.0360   1.0000
  14.500   1.1087   0.05588   0.04803   0.0150   0.0354   1.0000
  14.750   1.1084   0.05869   0.05093   0.0151   0.0351   1.0000
  15.000   1.1072   0.06166   0.05398   0.0151   0.0345   1.0000
  15.250   1.1061   0.06469   0.05710   0.0149   0.0342   1.0000
  15.500   1.1049   0.06780   0.06029   0.0147   0.0338   1.0000
  15.750   1.1035   0.07099   0.06354   0.0143   0.0333   1.0000
  16.000   1.1019   0.07422   0.06683   0.0139   0.0328   1.0000
  16.250   1.1011   0.07732   0.06996   0.0135   0.0324   1.0000
  16.500   1.1017   0.08024   0.07291   0.0132   0.0319   1.0000
  16.750   1.1039   0.08310   0.07587   0.0127   0.0316   1.0000
  17.000   1.1061   0.08598   0.07885   0.0122   0.0312   1.0000
  17.250   1.1081   0.08892   0.08189   0.0116   0.0308   1.0000
  17.500   1.1107   0.09173   0.08479   0.0110   0.0304   1.0000
<< Back to GOE 601 AIRFOIL (goe601-il)

Polar data table (+)

Polar graphs


<< Back to GOE 601 AIRFOIL (goe601-il)