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

GOE 508 AIRFOIL (goe508-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 508 AIRFOIL (goe508-il)
Reynolds number: 200,000
Max Cl/Cd: 71.34 at α=7.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe508-il-200000.txt
Download as CSV file: xf-goe508-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 508 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
  ------ -------- --------- --------- -------- -------- --------
 -10.750   0.0096   0.09943   0.09557  -0.1063   0.9366   0.0769
 -10.500   0.0274   0.09614   0.09225  -0.1100   0.9272   0.0790
 -10.250   0.0077   0.09052   0.08659  -0.1219   0.9150   0.0829
 -10.000   0.0386   0.08765   0.08367  -0.1212   0.9078   0.0836
  -9.750   0.0656   0.08508   0.08103  -0.1225   0.9007   0.0849
  -9.500   0.0778   0.08285   0.07877  -0.1231   0.8898   0.0864
  -9.250   0.0847   0.08023   0.07610  -0.1252   0.8804   0.0898
  -9.000   0.0391   0.07492   0.07077  -0.1319   0.8660   0.0928
  -8.750   0.0666   0.07342   0.06925  -0.1292   0.8591   0.0936
  -8.500   0.0851   0.07187   0.06766  -0.1280   0.8515   0.0950
  -8.250   0.0945   0.06989   0.06563  -0.1280   0.8442   0.0968
  -8.000   0.0923   0.06759   0.06334  -0.1280   0.8350   0.0991
  -7.750  -0.0308   0.06161   0.05708  -0.1328   0.8185   0.1033
  -7.500   0.0019   0.05798   0.05360  -0.1325   0.8137   0.1042
  -7.250  -0.0837   0.04086   0.03551  -0.1307   0.8058   0.0746
  -7.000  -0.0648   0.03993   0.03468  -0.1298   0.7986   0.0754
  -6.750  -0.0606   0.03622   0.03064  -0.1279   0.7927   0.0715
  -6.500  -0.0607   0.03235   0.02627  -0.1252   0.7867   0.0700
  -6.250  -0.0535   0.02972   0.02326  -0.1227   0.7795   0.0698
  -6.000  -0.0385   0.02750   0.02063  -0.1209   0.7742   0.0698
  -5.750  -0.0192   0.02580   0.01856  -0.1197   0.7695   0.0702
  -5.500  -0.0022   0.02459   0.01714  -0.1179   0.7624   0.0709
  -5.250   0.0200   0.02341   0.01568  -0.1168   0.7569   0.0720
  -5.000   0.0456   0.02241   0.01433  -0.1161   0.7526   0.0736
  -4.750   0.0663   0.02163   0.01350  -0.1149   0.7456   0.0752
  -4.500   0.0908   0.02113   0.01299  -0.1142   0.7397   0.0773
  -4.250   0.1185   0.02052   0.01226  -0.1139   0.7352   0.0797
  -4.000   0.1427   0.02004   0.01161  -0.1131   0.7298   0.0821
  -3.750   0.1659   0.01946   0.01106  -0.1122   0.7236   0.0846
  -3.500   0.1928   0.01910   0.01069  -0.1118   0.7188   0.0883
  -3.250   0.2220   0.01868   0.01010  -0.1118   0.7150   0.0931
  -3.000   0.2423   0.01841   0.00997  -0.1104   0.7081   0.0974
  -2.750   0.2679   0.01814   0.00961  -0.1097   0.7028   0.1034
  -2.500   0.2952   0.01770   0.00922  -0.1095   0.6988   0.1116
  -2.250   0.3197   0.01740   0.00896  -0.1088   0.6943   0.1216
  -2.000   0.3408   0.01717   0.00881  -0.1075   0.6885   0.1347
  -1.750   0.3658   0.01687   0.00857  -0.1069   0.6839   0.1524
  -1.500   0.3938   0.01661   0.00830  -0.1068   0.6801   0.1722
  -1.250   0.4165   0.01652   0.00829  -0.1058   0.6753   0.1899
  -1.000   0.4382   0.01643   0.00830  -0.1047   0.6701   0.2093
  -0.750   0.4631   0.01624   0.00819  -0.1041   0.6659   0.2361
  -0.500   0.4893   0.01588   0.00802  -0.1037   0.6624   0.2879
  -0.250   0.5051   0.01535   0.00809  -0.1016   0.6576   0.4303
   0.000   0.5162   0.01434   0.00840  -0.0976   0.6520   0.7867
   0.250   0.6369   0.01431   0.00836  -0.1147   0.6464   0.9638
   0.500   0.7076   0.01437   0.00820  -0.1229   0.6418   0.9858
   0.750   0.7665   0.01443   0.00821  -0.1295   0.6357   1.0000
   1.000   0.7864   0.01449   0.00818  -0.1281   0.6309   1.0000
   1.250   0.8104   0.01450   0.00805  -0.1273   0.6268   1.0000
   1.500   0.8277   0.01467   0.00818  -0.1255   0.6217   1.0000
   1.750   0.8449   0.01479   0.00828  -0.1235   0.6161   1.0000
   2.000   0.8681   0.01480   0.00818  -0.1226   0.6111   1.0000
   2.250   0.8919   0.01488   0.00815  -0.1217   0.6065   1.0000
   2.500   0.9059   0.01506   0.00835  -0.1192   0.6002   1.0000
   2.750   0.9278   0.01513   0.00835  -0.1180   0.5951   1.0000
   3.000   0.9542   0.01520   0.00831  -0.1176   0.5911   1.0000
   3.250   0.9705   0.01545   0.00858  -0.1155   0.5862   1.0000
   3.500   0.9886   0.01564   0.00877  -0.1137   0.5811   1.0000
   3.750   1.0124   0.01573   0.00879  -0.1129   0.5765   1.0000
   4.000   1.0403   0.01584   0.00878  -0.1129   0.5725   1.0000
   4.250   1.0521   0.01610   0.00914  -0.1099   0.5666   1.0000
   4.500   1.0727   0.01622   0.00923  -0.1085   0.5610   1.0000
   4.750   1.1002   0.01628   0.00918  -0.1083   0.5562   1.0000
   5.000   1.1143   0.01652   0.00948  -0.1058   0.5500   1.0000
   5.250   1.1331   0.01665   0.00961  -0.1041   0.5439   1.0000
   5.500   1.1603   0.01671   0.00957  -0.1039   0.5387   1.0000
   5.750   1.1729   0.01696   0.00988  -0.1011   0.5321   1.0000
   6.000   1.1911   0.01709   0.01001  -0.0993   0.5255   1.0000
   6.250   1.2174   0.01717   0.01000  -0.0989   0.5196   1.0000
   6.500   1.2248   0.01741   0.01034  -0.0952   0.5118   1.0000
   6.750   1.2459   0.01748   0.01035  -0.0939   0.5046   1.0000
   7.000   1.2555   0.01771   0.01063  -0.0905   0.4964   1.0000
   7.250   1.2705   0.01781   0.01069  -0.0881   0.4879   1.0000
   7.500   1.2751   0.01805   0.01096  -0.0839   0.4782   1.0000
   7.750   1.2875   0.01816   0.01100  -0.0810   0.4687   1.0000
   8.000   1.2834   0.01851   0.01143  -0.0755   0.4573   1.0000
   8.250   1.2878   0.01886   0.01176  -0.0716   0.4450   1.0000
   8.500   1.2914   0.01930   0.01216  -0.0678   0.4311   1.0000
   8.750   1.2919   0.01993   0.01274  -0.0637   0.4144   1.0000
   9.000   1.2920   0.02073   0.01348  -0.0600   0.3969   1.0000
   9.250   1.2926   0.02169   0.01433  -0.0565   0.3793   1.0000
   9.500   1.2946   0.02275   0.01528  -0.0535   0.3635   1.0000
   9.750   1.2979   0.02387   0.01628  -0.0508   0.3495   1.0000
  10.000   1.3033   0.02498   0.01729  -0.0485   0.3377   1.0000
  10.250   1.3115   0.02603   0.01832  -0.0466   0.3281   1.0000
  10.500   1.3217   0.02702   0.01917  -0.0450   0.3200   1.0000
  10.750   1.3310   0.02808   0.02027  -0.0434   0.3123   1.0000
  11.000   1.3426   0.02903   0.02114  -0.0420   0.3055   1.0000
  11.250   1.3546   0.03000   0.02212  -0.0407   0.2995   1.0000
  11.500   1.3657   0.03102   0.02316  -0.0394   0.2938   1.0000
  11.750   1.3813   0.03181   0.02384  -0.0385   0.2882   1.0000
  12.000   1.3930   0.03283   0.02492  -0.0373   0.2836   1.0000
  12.250   1.4043   0.03388   0.02604  -0.0362   0.2790   1.0000
  12.500   1.4186   0.03477   0.02692  -0.0353   0.2747   1.0000
  12.750   1.4414   0.03524   0.02729  -0.0349   0.2705   1.0000
  13.000   1.4479   0.03658   0.02877  -0.0336   0.2668   1.0000
  13.250   1.4582   0.03774   0.03001  -0.0326   0.2628   1.0000
  13.500   1.4705   0.03877   0.03106  -0.0316   0.2587   1.0000
  13.750   1.4971   0.03905   0.03121  -0.0316   0.2545   1.0000
  14.000   1.4971   0.04083   0.03318  -0.0300   0.2510   1.0000
  14.250   1.5032   0.04229   0.03475  -0.0289   0.2469   1.0000
  14.500   1.5130   0.04351   0.03599  -0.0281   0.2428   1.0000
  14.750   1.5368   0.04388   0.03625  -0.0278   0.2382   1.0000
  15.000   1.5330   0.04607   0.03864  -0.0264   0.2348   1.0000
  15.250   1.5362   0.04785   0.04054  -0.0254   0.2309   1.0000
  15.500   1.5432   0.04931   0.04203  -0.0247   0.2267   1.0000
  15.750   1.5621   0.04993   0.04256  -0.0242   0.2221   1.0000
  16.000   1.5529   0.05278   0.04563  -0.0231   0.2183   1.0000
  16.250   1.5516   0.05507   0.04803  -0.0224   0.2139   1.0000
  16.500   1.5576   0.05671   0.04968  -0.0218   0.2097   1.0000
  16.750   1.5668   0.05814   0.05112  -0.0213   0.2056   1.0000
  17.000   1.5595   0.06123   0.05441  -0.0208   0.2020   1.0000
  17.250   1.5583   0.06376   0.05705  -0.0205   0.1978   1.0000
  17.500   1.5641   0.06557   0.05886  -0.0202   0.1941   1.0000
  17.750   1.5669   0.06774   0.06109  -0.0199   0.1902   1.0000
  18.000   1.5582   0.07133   0.06488  -0.0199   0.1863   1.0000
  18.250   1.5553   0.07429   0.06794  -0.0200   0.1821   1.0000
  18.500   1.5617   0.07609   0.06969  -0.0199   0.1781   1.0000
  18.750   1.5515   0.08012   0.07392  -0.0203   0.1741   1.0000
  19.000   1.5427   0.08403   0.07798  -0.0208   0.1695   1.0000
  19.250   1.5428   0.08679   0.08077  -0.0212   0.1655   1.0000
<< Back to GOE 508 AIRFOIL (goe508-il)

Polar data table (+)

Polar graphs


<< Back to GOE 508 AIRFOIL (goe508-il)