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

GOE 288 AIRFOIL (goe288-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 288 AIRFOIL (goe288-il)
Reynolds number: 100,000
Max Cl/Cd: 50.63 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe288-il-100000-n5.txt
Download as CSV file: xf-goe288-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 288 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.2896   0.08822   0.08297  -0.0693   0.9809   0.0559
 -10.250  -0.4598   0.04972   0.04371  -0.1065   0.9445   0.0548
 -10.000  -0.4606   0.04295   0.03635  -0.1159   0.9268   0.0552
  -9.750  -0.4457   0.03925   0.03222  -0.1198   0.9133   0.0559
  -9.500  -0.4197   0.03730   0.03016  -0.1218   0.9028   0.0565
  -9.250  -0.3977   0.03555   0.02826  -0.1228   0.8885   0.0572
  -9.000  -0.3739   0.03386   0.02641  -0.1239   0.8749   0.0580
  -8.750  -0.3486   0.03224   0.02459  -0.1250   0.8623   0.0590
  -8.500  -0.3239   0.03076   0.02289  -0.1256   0.8484   0.0601
  -8.250  -0.2994   0.02942   0.02131  -0.1260   0.8341   0.0614
  -8.000  -0.2737   0.02827   0.01999  -0.1264   0.8211   0.0630
  -7.750  -0.2471   0.02739   0.01905  -0.1268   0.8087   0.0647
  -7.500  -0.2217   0.02654   0.01808  -0.1269   0.7947   0.0667
  -7.250  -0.1955   0.02566   0.01701  -0.1270   0.7820   0.0689
  -6.750  -0.1424   0.02420   0.01539  -0.1273   0.7586   0.0737
  -6.500  -0.1151   0.02353   0.01455  -0.1274   0.7488   0.0774
  -6.250  -0.0880   0.02295   0.01394  -0.1277   0.7386   0.0818
  -6.000  -0.0602   0.02239   0.01328  -0.1280   0.7299   0.0886
  -5.750  -0.0326   0.02189   0.01269  -0.1282   0.7205   0.0975
  -5.500  -0.0042   0.02141   0.01211  -0.1286   0.7128   0.1086
  -5.250   0.0236   0.02101   0.01165  -0.1289   0.7040   0.1212
  -5.000   0.0523   0.02066   0.01120  -0.1292   0.6972   0.1352
  -4.750   0.0806   0.02037   0.01084  -0.1295   0.6903   0.1492
  -4.500   0.1090   0.02010   0.01054  -0.1298   0.6834   0.1635
  -4.250   0.1378   0.01986   0.01025  -0.1301   0.6775   0.1798
  -4.000   0.1661   0.01968   0.01012  -0.1304   0.6709   0.1997
  -3.750   0.1944   0.01958   0.01008  -0.1307   0.6645   0.2216
  -3.500   0.2233   0.01953   0.01000  -0.1309   0.6593   0.2421
  -3.250   0.2520   0.01951   0.00996  -0.1310   0.6535   0.2603
  -3.000   0.2801   0.01952   0.00995  -0.1310   0.6462   0.2780
  -2.750   0.3089   0.01955   0.00987  -0.1309   0.6394   0.2957
  -2.500   0.3366   0.01962   0.00989  -0.1308   0.6312   0.3121
  -2.250   0.3646   0.01967   0.00983  -0.1306   0.6231   0.3272
  -2.000   0.3931   0.01974   0.00975  -0.1306   0.6162   0.3419
  -1.750   0.4207   0.01984   0.00982  -0.1304   0.6090   0.3549
  -1.500   0.4490   0.01991   0.00982  -0.1304   0.6033   0.3664
  -1.250   0.4779   0.02000   0.00976  -0.1304   0.5984   0.3788
  -1.000   0.5050   0.02009   0.00989  -0.1302   0.5920   0.3892
  -0.750   0.5329   0.02016   0.00990  -0.1302   0.5861   0.4002
  -0.500   0.5615   0.02022   0.00986  -0.1302   0.5811   0.4115
  -0.250   0.5884   0.02031   0.00998  -0.1300   0.5749   0.4218
   0.000   0.6159   0.02041   0.01003  -0.1299   0.5685   0.4346
   0.250   0.6434   0.02044   0.01005  -0.1297   0.5632   0.4455
   0.500   0.6708   0.02053   0.01011  -0.1296   0.5574   0.4558
   0.750   0.6972   0.02059   0.01021  -0.1293   0.5508   0.4639
   1.250   0.7518   0.02068   0.01023  -0.1289   0.5388   0.4803
   1.500   0.7778   0.02077   0.01034  -0.1286   0.5314   0.4894
   1.750   0.8049   0.02079   0.01034  -0.1284   0.5254   0.4981
   2.000   0.8309   0.02091   0.01047  -0.1280   0.5186   0.5077
   2.250   0.8568   0.02100   0.01061  -0.1277   0.5114   0.5175
   2.500   0.8841   0.02106   0.01059  -0.1275   0.5054   0.5298
   2.750   0.9085   0.02120   0.01085  -0.1269   0.4970   0.5419
   3.000   0.9342   0.02126   0.01093  -0.1265   0.4896   0.5565
   3.250   0.9589   0.02137   0.01110  -0.1259   0.4818   0.5747
   3.500   0.9831   0.02146   0.01129  -0.1253   0.4734   0.5988
   3.750   1.0073   0.02149   0.01142  -0.1246   0.4662   0.6345
   4.000   1.0278   0.02144   0.01170  -0.1232   0.4577   0.7082
   4.250   1.0487   0.02117   0.01164  -0.1214   0.4511   1.0000
   4.500   1.0717   0.02155   0.01202  -0.1208   0.4417   1.0000
   4.750   1.0961   0.02183   0.01216  -0.1202   0.4346   1.0000
   5.000   1.1184   0.02224   0.01256  -0.1195   0.4256   1.0000
   5.250   1.1415   0.02257   0.01278  -0.1188   0.4185   1.0000
   5.500   1.1634   0.02301   0.01320  -0.1179   0.4108   1.0000
   5.750   1.1854   0.02342   0.01355  -0.1171   0.4042   1.0000
   6.000   1.2076   0.02385   0.01390  -0.1163   0.3985   1.0000
   6.250   1.2283   0.02437   0.01443  -0.1154   0.3921   1.0000
   6.500   1.2496   0.02483   0.01483  -0.1145   0.3869   1.0000
   6.750   1.2710   0.02532   0.01525  -0.1136   0.3821   1.0000
   7.000   1.2901   0.02592   0.01588  -0.1125   0.3766   1.0000
   7.250   1.3098   0.02647   0.01641  -0.1114   0.3718   1.0000
   7.500   1.3315   0.02697   0.01682  -0.1106   0.3678   1.0000
   7.750   1.3499   0.02763   0.01751  -0.1095   0.3636   1.0000
   8.000   1.3667   0.02831   0.01825  -0.1080   0.3597   1.0000
   8.250   1.3845   0.02895   0.01891  -0.1068   0.3561   1.0000
   8.500   1.4045   0.02956   0.01949  -0.1058   0.3530   1.0000
   8.750   1.4290   0.03010   0.01996  -0.1055   0.3502   1.0000
   9.000   1.4418   0.03100   0.02097  -0.1038   0.3469   1.0000
   9.250   1.4556   0.03189   0.02195  -0.1022   0.3435   1.0000
   9.500   1.4713   0.03272   0.02283  -0.1009   0.3402   1.0000
   9.750   1.4893   0.03348   0.02362  -0.0998   0.3373   1.0000
  10.000   1.5109   0.03415   0.02428  -0.0993   0.3348   1.0000
  10.250   1.5357   0.03479   0.02489  -0.0992   0.3327   1.0000
  10.500   1.5424   0.03608   0.02636  -0.0970   0.3301   1.0000
  10.750   1.5506   0.03738   0.02782  -0.0952   0.3276   1.0000
  11.000   1.5605   0.03865   0.02921  -0.0936   0.3251   1.0000
  11.250   1.5725   0.03984   0.03049  -0.0923   0.3227   1.0000
  11.500   1.5873   0.04089   0.03160  -0.0913   0.3203   1.0000
  11.750   1.6072   0.04170   0.03245  -0.0907   0.3181   1.0000
  12.000   1.6328   0.04230   0.03304  -0.0907   0.3162   1.0000
  12.250   1.6319   0.04427   0.03520  -0.0885   0.3140   1.0000
  12.500   1.6188   0.04711   0.03829  -0.0858   0.3115   1.0000
  12.750   1.6062   0.05020   0.04159  -0.0836   0.3088   1.0000
  13.000   1.5983   0.05312   0.04469  -0.0819   0.3060   1.0000
  13.250   1.6025   0.05513   0.04680  -0.0809   0.3034   1.0000
  13.500   1.6251   0.05553   0.04722  -0.0804   0.3008   1.0000
  13.750   1.6668   0.05449   0.04609  -0.0807   0.2983   1.0000
  14.000   1.3207   0.09936   0.09192  -0.0831   0.2825   1.0000
  14.250   1.3528   0.09767   0.09028  -0.0819   0.2820   1.0000
  14.500   1.3963   0.09430   0.08692  -0.0803   0.2816   1.0000
  14.750   1.4536   0.08911   0.08174  -0.0785   0.2814   1.0000
<< Back to GOE 288 AIRFOIL (goe288-il)

Polar data table (+)

Polar graphs


<< Back to GOE 288 AIRFOIL (goe288-il)