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=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 304 (FRIEDRICHSHAFEN G02) AIRFOIL (goe304-il)
Reynolds number: 500,000
Max Cl/Cd: 109.66 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe304-il-500000-n5.txt
Download as CSV file: xf-goe304-il-500000-n5.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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.1384   0.09049   0.08797  -0.0623   0.9349   0.0236
  -8.250  -0.1318   0.08716   0.08459  -0.0652   0.9176   0.0245
  -8.000  -0.1249   0.08421   0.08160  -0.0671   0.9005   0.0245
  -7.750  -0.1179   0.08129   0.07861  -0.0692   0.8837   0.0246
  -7.500  -0.1085   0.07823   0.07550  -0.0723   0.8685   0.0246
  -7.250  -0.0949   0.07479   0.07199  -0.0766   0.8550   0.0246
  -7.000  -0.0814   0.07209   0.06923  -0.0778   0.8432   0.0248
  -6.750  -0.0668   0.07009   0.06719  -0.0784   0.8313   0.0250
  -6.500  -0.0500   0.06795   0.06500  -0.0803   0.8200   0.0255
  -6.250  -0.0316   0.06552   0.06251  -0.0833   0.8092   0.0262
  -6.000  -0.0079   0.06209   0.05900  -0.0890   0.7980   0.0279
  -5.750   0.0205   0.05781   0.05460  -0.0968   0.7878   0.0282
  -5.250   0.0746   0.05048   0.04706  -0.1071   0.7680   0.0282
  -5.000   0.1041   0.04658   0.04301  -0.1122   0.7586   0.0283
  -4.750   0.1266   0.04443   0.04080  -0.1136   0.7485   0.0281
  -4.500   0.1470   0.04295   0.03927  -0.1141   0.7393   0.0276
  -4.250   0.1753   0.04007   0.03628  -0.1172   0.7299   0.0270
  -4.000   0.2059   0.03693   0.03298  -0.1207   0.7217   0.0267
  -3.750   0.2370   0.03385   0.02973  -0.1237   0.7133   0.0266
  -3.500   0.2679   0.03089   0.02657  -0.1263   0.7057   0.0267
  -3.250   0.2996   0.02780   0.02326  -0.1285   0.6980   0.0273
  -2.750   0.3621   0.02063   0.01540  -0.1318   0.6839   0.0276
  -2.500   0.3928   0.01514   0.00908  -0.1329   0.6774   0.0286
  -2.250   0.4203   0.01478   0.00862  -0.1329   0.6703   0.0289
  -2.000   0.4480   0.01434   0.00807  -0.1329   0.6629   0.0294
  -1.500   0.5039   0.01273   0.00605  -0.1328   0.6492   0.0305
  -1.250   0.5317   0.01205   0.00517  -0.1327   0.6423   0.0311
  -1.000   0.5595   0.01152   0.00448  -0.1326   0.6360   0.0318
  -0.750   0.5873   0.01115   0.00399  -0.1324   0.6288   0.0324
  -0.250   0.6426   0.01084   0.00364  -0.1322   0.6157   0.0337
   0.000   0.6701   0.01073   0.00350  -0.1320   0.6085   0.0347
   0.250   0.6976   0.01058   0.00330  -0.1318   0.6018   0.0356
   0.500   0.7253   0.01042   0.00310  -0.1316   0.5944   0.0365
   0.750   0.7523   0.01032   0.00296  -0.1314   0.5851   0.0373
   1.000   0.7788   0.01030   0.00290  -0.1310   0.5698   0.0383
   1.250   0.8051   0.01031   0.00285  -0.1306   0.5503   0.0395
   1.500   0.8314   0.01036   0.00282  -0.1302   0.5332   0.0413
   1.750   0.8580   0.01036   0.00279  -0.1299   0.5205   0.0430
   2.000   0.8851   0.01037   0.00280  -0.1297   0.5099   0.0448
   2.250   0.9118   0.01042   0.00282  -0.1294   0.4988   0.0469
   2.500   0.9380   0.01049   0.00285  -0.1290   0.4850   0.0492
   2.750   0.9639   0.01061   0.00292  -0.1286   0.4693   0.0526
   3.000   0.9896   0.01074   0.00299  -0.1282   0.4530   0.0561
   3.250   1.0154   0.01088   0.00309  -0.1278   0.4382   0.0604
   3.500   1.0413   0.01100   0.00320  -0.1274   0.4252   0.0678
   3.750   1.0671   0.01112   0.00336  -0.1270   0.4125   0.1050
   4.250   1.1141   0.01016   0.00391  -0.1258   0.3814   1.0000
   4.500   1.1389   0.01044   0.00409  -0.1252   0.3656   1.0000
   4.750   1.1632   0.01074   0.00429  -0.1246   0.3487   1.0000
   5.000   1.1870   0.01108   0.00453  -0.1239   0.3289   1.0000
   5.250   1.2101   0.01147   0.00480  -0.1231   0.3074   1.0000
   5.500   1.2329   0.01187   0.00508  -0.1223   0.2876   1.0000
   5.750   1.2556   0.01227   0.00538  -0.1214   0.2692   1.0000
   6.000   1.2781   0.01268   0.00570  -0.1206   0.2527   1.0000
   6.250   1.3004   0.01308   0.00602  -0.1197   0.2384   1.0000
   6.500   1.3223   0.01351   0.00638  -0.1187   0.2245   1.0000
   6.750   1.3439   0.01393   0.00674  -0.1177   0.2114   1.0000
   7.000   1.3659   0.01430   0.00708  -0.1168   0.2019   1.0000
   7.250   1.3871   0.01472   0.00745  -0.1158   0.1927   1.0000
   7.500   1.4085   0.01510   0.00782  -0.1148   0.1842   1.0000
   7.750   1.4290   0.01552   0.00822  -0.1136   0.1764   1.0000
   8.000   1.4497   0.01591   0.00860  -0.1125   0.1691   1.0000
   8.250   1.4687   0.01637   0.00905  -0.1111   0.1615   1.0000
   8.500   1.4866   0.01682   0.00948  -0.1095   0.1527   1.0000
   8.750   1.5030   0.01732   0.00996  -0.1077   0.1438   1.0000
   9.000   1.5183   0.01791   0.01050  -0.1058   0.1334   1.0000
   9.250   1.5333   0.01851   0.01109  -0.1039   0.1225   1.0000
   9.500   1.5472   0.01920   0.01173  -0.1019   0.1119   1.0000
   9.750   1.5605   0.01993   0.01244  -0.0999   0.1038   1.0000
  10.000   1.5731   0.02072   0.01320  -0.0979   0.0975   1.0000
  10.250   1.5868   0.02146   0.01397  -0.0961   0.0936   1.0000
  10.500   1.5997   0.02227   0.01480  -0.0942   0.0903   1.0000
  10.750   1.6113   0.02318   0.01573  -0.0923   0.0871   1.0000
  11.000   1.6235   0.02409   0.01668  -0.0906   0.0844   1.0000
  11.250   1.6369   0.02493   0.01759  -0.0891   0.0825   1.0000
  11.500   1.6489   0.02590   0.01861  -0.0875   0.0803   1.0000
  11.750   1.6596   0.02700   0.01975  -0.0859   0.0780   1.0000
  12.000   1.6687   0.02825   0.02104  -0.0842   0.0755   1.0000
  12.250   1.6775   0.02956   0.02240  -0.0826   0.0733   1.0000
  12.500   1.6904   0.03058   0.02352  -0.0814   0.0717   1.0000
  12.750   1.7014   0.03177   0.02478  -0.0802   0.0692   1.0000
  13.000   1.7102   0.03318   0.02624  -0.0789   0.0662   1.0000
  13.250   1.7170   0.03479   0.02790  -0.0775   0.0632   1.0000
  13.500   1.7266   0.03620   0.02938  -0.0764   0.0591   1.0000
  13.750   1.7319   0.03802   0.03121  -0.0752   0.0523   1.0000
  14.000   1.7320   0.04038   0.03351  -0.0738   0.0427   1.0000
  14.250   1.7308   0.04295   0.03608  -0.0725   0.0372   1.0000
  14.500   1.7302   0.04554   0.03872  -0.0714   0.0338   1.0000
  14.750   1.7297   0.04819   0.04143  -0.0704   0.0315   1.0000
  15.000   1.7285   0.05100   0.04432  -0.0696   0.0296   1.0000
  15.250   1.7286   0.05376   0.04716  -0.0690   0.0280   1.0000
  15.500   1.7267   0.05679   0.05027  -0.0684   0.0266   1.0000
  16.000   1.7220   0.06318   0.05685  -0.0677   0.0243   1.0000
  16.250   1.7194   0.06651   0.06028  -0.0675   0.0232   1.0000
  16.500   1.7151   0.07012   0.06399  -0.0675   0.0223   1.0000
  16.750   1.7087   0.07406   0.06802  -0.0676   0.0215   1.0000
  17.000   1.7038   0.07791   0.07198  -0.0679   0.0208   1.0000
  17.250   1.6989   0.08182   0.07600  -0.0683   0.0202   1.0000
  17.500   1.6927   0.08597   0.08026  -0.0688   0.0195   1.0000
  17.750   1.6850   0.09038   0.08478  -0.0695   0.0189   1.0000
  18.000   1.6759   0.09509   0.08959  -0.0704   0.0184   1.0000
  18.250   1.6656   0.10008   0.09468  -0.0715   0.0179   1.0000
  18.500   1.6554   0.10510   0.09982  -0.0728   0.0175   1.0000
  18.750   1.6466   0.10995   0.10479  -0.0742   0.0171   1.0000
  19.000   1.6371   0.11499   0.10995  -0.0757   0.0167   1.0000
  19.250   1.6270   0.12020   0.11528  -0.0775   0.0163   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)