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

GOE 304 (FRIEDRICHSHAFEN G02) AIRFOIL (goe304-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 304 (FRIEDRICHSHAFEN G02) AIRFOIL (goe304-il)
Reynolds number: 50,000
Max Cl/Cd: 33.06 at α=9°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe304-il-50000.txt
Download as CSV file: xf-goe304-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 304 (FRIEDRICHSHAFEN G02) 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.2628   0.12722   0.11995  -0.0357   1.0000   0.1109
  -9.500  -0.2750   0.12822   0.12110  -0.0358   1.0000   0.1118
  -9.250  -0.2917   0.12957   0.12261  -0.0350   1.0000   0.1122
  -9.000  -0.2579   0.11935   0.11233  -0.0334   1.0000   0.1160
  -8.750  -0.2565   0.11704   0.11010  -0.0320   1.0000   0.1190
  -8.500  -0.2607   0.11564   0.10879  -0.0304   1.0000   0.1217
  -8.250  -0.2701   0.11497   0.10825  -0.0286   1.0000   0.1241
  -8.000  -0.2862   0.11521   0.10863  -0.0262   1.0000   0.1256
  -7.750  -0.3074   0.11599   0.10955  -0.0232   1.0000   0.1264
  -7.500  -0.3280   0.11686   0.11057  -0.0213   1.0000   0.1269
  -7.250  -0.3454   0.11779   0.11163  -0.0216   1.0000   0.1274
  -7.000  -0.3363   0.11180   0.10569  -0.0175   1.0000   0.1292
  -6.750  -0.3342   0.10864   0.10258  -0.0146   1.0000   0.1318
  -6.500  -0.3392   0.10695   0.10097  -0.0127   1.0000   0.1344
  -6.250  -0.3457   0.10560   0.09969  -0.0116   1.0000   0.1371
  -6.000  -0.3524   0.10458   0.09875  -0.0118   1.0000   0.1402
  -5.750  -0.3577   0.10522   0.09941  -0.0169   1.0000   0.1430
  -5.500  -0.3572   0.10232   0.09658  -0.0169   1.0000   0.1445
  -5.250  -0.3565   0.09861   0.09293  -0.0125   1.0000   0.1471
  -5.000  -0.3546   0.09631   0.09066  -0.0113   1.0000   0.1514
  -4.750  -0.3446   0.09578   0.09009  -0.0180   1.0000   0.1591
  -4.500  -0.3404   0.09248   0.08685  -0.0177   1.0000   0.1619
  -4.250  -0.3381   0.08962   0.08404  -0.0148   1.0000   0.1665
  -4.000  -0.3138   0.08885   0.08314  -0.0246   1.0000   0.1772
  -3.750  -0.3146   0.08503   0.07943  -0.0200   1.0000   0.1802
  -3.500  -0.3047   0.08276   0.07716  -0.0203   1.0000   0.1881
  -3.250  -0.2866   0.08035   0.07470  -0.0243   1.0000   0.1966
  -3.000  -0.2614   0.07886   0.07311  -0.0301   1.0000   0.2109
  -2.750  -0.2586   0.07569   0.07004  -0.0268   1.0000   0.2163
  -2.500  -0.2248   0.07303   0.06733  -0.0321   0.9947   0.2342
  -2.250  -0.1792   0.07039   0.06459  -0.0401   0.9852   0.2636
  -2.000  -0.1456   0.06792   0.06209  -0.0442   0.9763   0.2974
  -1.750  -0.1174   0.06537   0.05957  -0.0460   0.9680   0.3353
  -1.500  -0.0938   0.06303   0.05728  -0.0467   0.9594   0.3863
  -0.750  -0.0572   0.05607   0.05064  -0.0368   0.9357   0.5761
  -0.500  -0.0377   0.05373   0.04838  -0.0347   0.9278   0.6315
  -0.250  -0.0152   0.05192   0.04659  -0.0349   0.9192   0.6745
   0.000   0.0167   0.04981   0.04452  -0.0359   0.9108   0.7134
   0.250   0.0599   0.04857   0.04321  -0.0422   0.9006   0.7294
   0.500   0.3229   0.05217   0.04443  -0.1061   0.8811   0.3216
   0.750   0.3800   0.05233   0.04368  -0.1119   0.8709   0.2433
   1.000   0.4292   0.05138   0.04242  -0.1158   0.8618   0.2173
   1.250   0.4571   0.05146   0.04218  -0.1165   0.8510   0.2025
   1.500   0.5085   0.05120   0.04146  -0.1201   0.8422   0.1908
   1.750   0.5302   0.05153   0.04163  -0.1199   0.8308   0.1906
   2.000   0.5603   0.05187   0.04175  -0.1207   0.8205   0.1910
   2.250   0.6050   0.05179   0.04136  -0.1229   0.8107   0.1912
   2.500   0.6206   0.05263   0.04206  -0.1217   0.7987   0.1939
   2.750   0.6523   0.05288   0.04230  -0.1224   0.7879   0.2039
   3.000   0.6913   0.05289   0.04225  -0.1236   0.7772   0.2149
   3.250   0.7071   0.05382   0.04321  -0.1225   0.7643   0.2286
   3.500   0.7347   0.05428   0.04381  -0.1227   0.7519   0.2580
   3.750   0.7837   0.05233   0.04311  -0.1238   0.7419   1.0000
   4.000   0.7973   0.05368   0.04410  -0.1220   0.7273   1.0000
   4.250   0.8151   0.05486   0.04508  -0.1207   0.7126   1.0000
   4.500   0.8363   0.05589   0.04598  -0.1198   0.6980   1.0000
   4.750   0.8608   0.05672   0.04671  -0.1190   0.6837   1.0000
   5.000   0.8912   0.05713   0.04706  -0.1186   0.6699   1.0000
   5.250   0.9390   0.05628   0.04616  -0.1190   0.6580   1.0000
   5.500   0.9551   0.05738   0.04727  -0.1175   0.6423   1.0000
   5.750   0.9726   0.05839   0.04829  -0.1159   0.6267   1.0000
   6.000   0.9900   0.05944   0.04936  -0.1144   0.6111   1.0000
   6.250   1.0092   0.06032   0.05029  -0.1129   0.5956   1.0000
   6.500   1.0292   0.06110   0.05111  -0.1114   0.5803   1.0000
   6.750   1.0507   0.06172   0.05177  -0.1099   0.5653   1.0000
   7.000   1.1249   0.05678   0.04693  -0.1091   0.5579   1.0000
   7.250   1.1408   0.05763   0.04785  -0.1071   0.5421   1.0000
   7.500   1.1570   0.05843   0.04872  -0.1050   0.5265   1.0000
   7.750   1.1724   0.05932   0.04969  -0.1030   0.5111   1.0000
   8.000   1.1879   0.06020   0.05065  -0.1009   0.4960   1.0000
   8.250   1.2324   0.05755   0.04808  -0.0991   0.4820   1.0000
   8.500   1.3645   0.04638   0.03692  -0.1015   0.4624   1.0000
   8.750   1.4237   0.04335   0.03377  -0.1021   0.4422   1.0000
   9.000   1.4458   0.04373   0.03415  -0.1005   0.4243   1.0000
   9.250   1.4641   0.04453   0.03499  -0.0986   0.4072   1.0000
   9.500   1.4826   0.04541   0.03593  -0.0968   0.3905   1.0000
   9.750   1.5002   0.04639   0.03694  -0.0950   0.3739   1.0000
  10.000   1.5172   0.04745   0.03803  -0.0931   0.3576   1.0000
  10.250   1.5342   0.04861   0.03920  -0.0913   0.3418   1.0000
  10.500   1.5542   0.04987   0.04048  -0.0899   0.3268   1.0000
  10.750   1.5479   0.05267   0.04353  -0.0863   0.3160   1.0000
  11.000   1.5471   0.05544   0.04646  -0.0834   0.3060   1.0000
  11.250   1.5701   0.05692   0.04794  -0.0825   0.2938   1.0000
  11.500   1.5325   0.06153   0.05287  -0.0768   0.2897   1.0000
  11.750   1.3338   0.08442   0.07608  -0.0733   0.2996   1.0000
  12.000   1.0956   0.13181   0.12309  -0.0941   0.3032   1.0000
<< Back to GOE 304 (FRIEDRICHSHAFEN G02) AIRFOIL (goe304-il)

Polar data table (+)

Polar graphs


<< Back to GOE 304 (FRIEDRICHSHAFEN G02) AIRFOIL (goe304-il)