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

GOE 512 AIRFOIL (goe512-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 512 AIRFOIL (goe512-il)
Reynolds number: 100,000
Max Cl/Cd: 48.81 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe512-il-100000.txt
Download as CSV file: xf-goe512-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 512 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3135   0.09865   0.09503  -0.0358   0.9792   0.1265
  -7.500  -0.3885   0.10215   0.09832  -0.0155   1.0000   0.1188
  -7.250  -0.4018   0.10022   0.09645  -0.0161   0.9975   0.1222
  -7.000  -0.4134   0.09516   0.09131  -0.0301   0.9829   0.1276
  -6.750  -0.2321   0.07926   0.07560  -0.0435   0.9487   0.1487
  -6.500  -0.2469   0.07594   0.07225  -0.0464   0.9366   0.1557
  -6.250  -0.3351   0.08235   0.07841  -0.0393   0.9566   0.1466
  -6.000  -0.3286   0.07839   0.07421  -0.0465   0.9420   0.1586
  -5.750  -0.3077   0.07531   0.07127  -0.0435   0.9336   0.1628
  -5.500  -0.2932   0.07163   0.06741  -0.0475   0.9228   0.1755
  -5.250  -0.2686   0.06846   0.06425  -0.0476   0.9147   0.1814
  -5.000  -0.2948   0.04790   0.04209  -0.0509   0.9020   0.0962
  -4.750  -0.2695   0.04298   0.03662  -0.0515   0.8965   0.0949
  -4.500  -0.2581   0.03982   0.03293  -0.0489   0.8853   0.0952
  -4.250  -0.2234   0.03632   0.02868  -0.0497   0.8807   0.0962
  -4.000  -0.2069   0.03455   0.02643  -0.0474   0.8692   0.0990
  -3.750  -0.1628   0.03252   0.02427  -0.0500   0.8651   0.1041
  -3.500  -0.1414   0.03153   0.02300  -0.0484   0.8532   0.1092
  -3.250  -0.0932   0.02966   0.02093  -0.0514   0.8495   0.1162
  -3.000  -0.0702   0.02898   0.02013  -0.0501   0.8375   0.1231
  -2.750  -0.0195   0.02740   0.01852  -0.0534   0.8338   0.1348
  -2.500   0.0046   0.02671   0.01780  -0.0523   0.8219   0.1458
  -2.250   0.0555   0.02528   0.01648  -0.0556   0.8182   0.1690
  -2.000   0.0804   0.02484   0.01623  -0.0544   0.8064   0.1955
  -1.750   0.1241   0.02453   0.01597  -0.0557   0.8019   0.2604
  -1.500   0.1441   0.02439   0.01573  -0.0536   0.7894   0.3018
  -1.250   0.1988   0.02314   0.01440  -0.0578   0.7856   0.3450
  -1.000   0.2320   0.02234   0.01371  -0.0584   0.7736   0.3698
  -0.750   0.3032   0.02088   0.01233  -0.0661   0.7688   0.4095
  -0.500   0.3479   0.02018   0.01175  -0.0690   0.7558   0.4391
  -0.250   0.4008   0.01939   0.01102  -0.0734   0.7430   0.4701
   0.000   0.4468   0.01863   0.01031  -0.0763   0.7285   0.5004
   0.250   0.6925   0.01748   0.00993  -0.1180   0.6992   0.9889
   0.500   0.7504   0.01745   0.00963  -0.1240   0.6784   1.0000
   0.750   0.7750   0.01763   0.00959  -0.1231   0.6616   1.0000
   1.000   0.7989   0.01785   0.00960  -0.1221   0.6458   1.0000
   1.250   0.8222   0.01809   0.00964  -0.1211   0.6309   1.0000
   1.500   0.8466   0.01833   0.00968  -0.1202   0.6173   1.0000
   1.750   0.8660   0.01861   0.00984  -0.1185   0.6034   1.0000
   2.000   0.8842   0.01893   0.01007  -0.1166   0.5905   1.0000
   2.250   0.9045   0.01922   0.01025  -0.1150   0.5785   1.0000
   2.500   0.9282   0.01947   0.01033  -0.1141   0.5675   1.0000
   2.750   0.9430   0.01980   0.01064  -0.1116   0.5554   1.0000
   3.000   0.9615   0.02010   0.01086  -0.1098   0.5443   1.0000
   3.250   0.9848   0.02033   0.01093  -0.1088   0.5337   1.0000
   3.500   0.9981   0.02066   0.01126  -0.1059   0.5217   1.0000
   3.750   1.0153   0.02098   0.01152  -0.1039   0.5107   1.0000
   4.000   1.0378   0.02126   0.01163  -0.1027   0.5002   1.0000
   4.250   1.0518   0.02163   0.01201  -0.1001   0.4886   1.0000
   4.500   1.0688   0.02204   0.01237  -0.0980   0.4778   1.0000
   4.750   1.0932   0.02242   0.01256  -0.0974   0.4677   1.0000
   5.000   1.1056   0.02291   0.01311  -0.0945   0.4566   1.0000
   5.250   1.1242   0.02343   0.01357  -0.0928   0.4465   1.0000
   5.500   1.1461   0.02392   0.01394  -0.0917   0.4367   1.0000
   5.750   1.1585   0.02453   0.01460  -0.0890   0.4270   1.0000
   6.000   1.1859   0.02506   0.01496  -0.0890   0.4185   1.0000
   6.250   1.1933   0.02573   0.01577  -0.0854   0.4098   1.0000
   6.500   1.2189   0.02630   0.01622  -0.0851   0.4027   1.0000
   6.750   1.2294   0.02706   0.01711  -0.0821   0.3957   1.0000
   7.000   1.2461   0.02771   0.01779  -0.0803   0.3893   1.0000
   7.250   1.2734   0.02839   0.01836  -0.0804   0.3837   1.0000
   7.500   1.2768   0.02917   0.01935  -0.0762   0.3777   1.0000
   7.750   1.2942   0.02981   0.02003  -0.0745   0.3722   1.0000
   8.000   1.3239   0.03051   0.02062  -0.0752   0.3676   1.0000
   8.250   1.3230   0.03146   0.02183  -0.0703   0.3628   1.0000
   8.500   1.3343   0.03220   0.02267  -0.0676   0.3578   1.0000
   8.750   1.3611   0.03269   0.02309  -0.0677   0.3530   1.0000
   9.000   1.3661   0.03363   0.02418  -0.0640   0.3483   1.0000
   9.250   1.3687   0.03453   0.02525  -0.0600   0.3436   1.0000
   9.500   1.3871   0.03508   0.02581  -0.0586   0.3390   1.0000
   9.750   1.4131   0.03574   0.02644  -0.0586   0.3347   1.0000
  10.000   1.4022   0.03700   0.02797  -0.0525   0.3309   1.0000
  10.250   1.4038   0.03800   0.02911  -0.0485   0.3267   1.0000
  10.500   1.4239   0.03862   0.02976  -0.0476   0.3227   1.0000
  10.750   1.4487   0.03937   0.03051  -0.0475   0.3188   1.0000
  11.000   1.4237   0.04102   0.03242  -0.0395   0.3158   1.0000
  11.250   1.4087   0.04248   0.03407  -0.0333   0.3125   1.0000
  11.500   1.4139   0.04353   0.03521  -0.0304   0.3089   1.0000
  11.750   1.4701   0.04360   0.03519  -0.0350   0.3044   1.0000
  12.000   1.4291   0.04590   0.03774  -0.0256   0.3019   1.0000
  12.250   1.3795   0.04921   0.04132  -0.0169   0.2992   1.0000
  12.500   1.3315   0.05335   0.04567  -0.0102   0.2960   1.0000
  12.750   1.3807   0.05222   0.04452  -0.0118   0.2915   1.0000
  13.000   1.4044   0.05283   0.04517  -0.0115   0.2874   1.0000
  13.750   1.4301   0.05590   0.04843  -0.0065   0.2732   1.0000
  14.000   0.7710   0.15772   0.15046  -0.0296   0.3046   1.0000
  14.250   0.7324   0.16107   0.15380  -0.0319   0.2977   1.0000
<< Back to GOE 512 AIRFOIL (goe512-il)

Polar data table (+)

Polar graphs


<< Back to GOE 512 AIRFOIL (goe512-il)