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

GOE 331 (PFALZ 60) AIRFOIL (goe331-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 331 (PFALZ 60) AIRFOIL (goe331-il)
Reynolds number: 1,000,000
Max Cl/Cd: 119.89 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe331-il-1000000.txt
Download as CSV file: xf-goe331-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 331 (PFALZ 60) AIRFOIL                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.1172   0.10371   0.10192  -0.0644   0.9322   0.0152
  -9.750  -0.1116   0.10113   0.09927  -0.0648   0.9081   0.0157
  -9.500  -0.1059   0.09845   0.09646  -0.0656   0.8767   0.0160
  -9.250  -0.0996   0.09570   0.09358  -0.0672   0.8533   0.0160
  -9.000  -0.0943   0.09290   0.09069  -0.0698   0.8375   0.0161
  -8.750  -0.0887   0.09001   0.08773  -0.0725   0.8259   0.0161
  -8.500  -0.0798   0.08673   0.08441  -0.0731   0.8169   0.0162
  -8.250  -0.0697   0.08402   0.08166  -0.0734   0.8090   0.0163
  -8.000  -0.0598   0.08152   0.07914  -0.0745   0.8013   0.0164
  -7.750  -0.0499   0.07911   0.07669  -0.0759   0.7944   0.0165
  -7.500  -0.0400   0.07670   0.07427  -0.0777   0.7883   0.0167
  -7.250  -0.0271   0.07411   0.07165  -0.0805   0.7822   0.0169
  -7.000  -0.0117   0.07134   0.06886  -0.0839   0.7766   0.0171
  -6.750   0.0054   0.06843   0.06593  -0.0877   0.7713   0.0175
  -6.500   0.0245   0.06540   0.06286  -0.0918   0.7659   0.0182
  -6.250   0.0534   0.06141   0.05879  -0.1001   0.7607   0.0188
  -6.000   0.0834   0.05734   0.05465  -0.1079   0.7561   0.0189
  -5.750   0.1109   0.05367   0.05089  -0.1129   0.7508   0.0189
  -5.500   0.1316   0.05026   0.04740  -0.1155   0.7455   0.0190
  -5.250   0.1500   0.04786   0.04499  -0.1163   0.7407   0.0191
  -5.000   0.1716   0.04560   0.04271  -0.1179   0.7356   0.0192
  -4.750   0.1954   0.04340   0.04044  -0.1199   0.7307   0.0194
  -4.500   0.2213   0.04113   0.03809  -0.1221   0.7257   0.0197
  -4.250   0.2487   0.03878   0.03568  -0.1244   0.7202   0.0201
  -4.000   0.2769   0.03645   0.03323  -0.1264   0.7143   0.0208
  -3.750   0.3150   0.03363   0.03018  -0.1291   0.7091   0.0217
  -3.500   0.3461   0.03118   0.02757  -0.1306   0.7036   0.0218
  -3.250   0.3752   0.02897   0.02519  -0.1315   0.6977   0.0218
  -3.000   0.3987   0.02646   0.02259  -0.1324   0.6919   0.0220
  -2.750   0.4241   0.02498   0.02105  -0.1331   0.6848   0.0222
  -2.500   0.4500   0.02368   0.01963  -0.1335   0.6770   0.0224
  -2.250   0.4773   0.02237   0.01823  -0.1340   0.6692   0.0227
  -2.000   0.5049   0.02110   0.01683  -0.1342   0.6618   0.0233
  -1.750   0.5378   0.01972   0.01518  -0.1339   0.6550   0.0250
  -1.500   0.5668   0.01856   0.01375  -0.1337   0.6469   0.0251
  -1.250   0.5937   0.01671   0.01176  -0.1340   0.6387   0.0253
  -0.750   0.6470   0.01510   0.00998  -0.1344   0.6174   0.0259
  -0.500   0.6742   0.01444   0.00921  -0.1344   0.6062   0.0263
  -0.250   0.7015   0.01380   0.00842  -0.1342   0.5943   0.0270
   0.000   0.1236   0.00831   0.00512  -0.0073   0.5728   0.1270
   0.250   0.7575   0.01237   0.00655  -0.1336   0.5685   0.0293
   0.500   0.7840   0.01196   0.00607  -0.1335   0.5545   0.0299
   0.750   0.8106   0.01167   0.00566  -0.1333   0.5393   0.0312
   1.000   0.8376   0.01142   0.00517  -0.1329   0.5225   0.0335
   1.250   0.8634   0.01111   0.00480  -0.1326   0.5039   0.0343
   1.500   0.8889   0.01101   0.00459  -0.1323   0.4830   0.0357
   3.750   1.1125   0.01123   0.00387  -0.1276   0.3157   0.0349
   4.000   1.1383   0.01132   0.00394  -0.1273   0.3107   0.0359
   4.250   1.1634   0.01145   0.00405  -0.1269   0.3055   0.0357
   4.500   1.1888   0.01157   0.00416  -0.1265   0.3014   0.0361
   4.750   1.2149   0.01165   0.00426  -0.1263   0.2984   0.0371
   5.000   1.2402   0.01177   0.00438  -0.1260   0.2950   0.0386
   5.250   1.2649   0.01195   0.00453  -0.1255   0.2911   0.0413
   5.500   1.2887   0.01206   0.00485  -0.1250   0.2862   0.2423
   5.750   1.3116   0.01099   0.00519  -0.1245   0.2840   1.0000
   6.000   1.3368   0.01115   0.00535  -0.1242   0.2809   1.0000
   6.250   1.3612   0.01136   0.00554  -0.1237   0.2779   1.0000
   6.500   1.3847   0.01160   0.00576  -0.1231   0.2746   1.0000
   6.750   1.4068   0.01191   0.00605  -0.1223   0.2707   1.0000
   7.000   1.4310   0.01210   0.00625  -0.1218   0.2688   1.0000
   7.250   1.4551   0.01227   0.00646  -0.1213   0.2667   1.0000
   7.500   1.4784   0.01247   0.00667  -0.1207   0.2641   1.0000
   7.750   1.5005   0.01271   0.00692  -0.1199   0.2616   1.0000
   8.000   1.5206   0.01299   0.00720  -0.1187   0.2584   1.0000
   8.250   1.5392   0.01334   0.00754  -0.1172   0.2539   1.0000
   8.500   1.5623   0.01350   0.00774  -0.1166   0.2511   1.0000
   8.750   1.5836   0.01374   0.00800  -0.1157   0.2478   1.0000
   9.000   1.6035   0.01405   0.00832  -0.1146   0.2442   1.0000
   9.250   1.6216   0.01445   0.00872  -0.1132   0.2405   1.0000
   9.500   1.6421   0.01474   0.00905  -0.1123   0.2375   1.0000
   9.750   1.6631   0.01501   0.00935  -0.1115   0.2340   1.0000
  10.000   1.6821   0.01538   0.00974  -0.1103   0.2295   1.0000
  10.250   1.6988   0.01587   0.01021  -0.1089   0.2238   1.0000
  10.500   1.7191   0.01618   0.01057  -0.1080   0.2191   1.0000
  10.750   1.7360   0.01669   0.01107  -0.1067   0.2127   1.0000
  11.000   1.7530   0.01719   0.01159  -0.1054   0.2069   1.0000
  11.250   1.7693   0.01774   0.01215  -0.1041   0.1997   1.0000
  11.500   1.7838   0.01841   0.01280  -0.1026   0.1909   1.0000
  11.750   1.7949   0.01931   0.01364  -0.1007   0.1768   1.0000
  12.000   1.7993   0.02066   0.01486  -0.0981   0.1557   1.0000
  12.250   1.7989   0.02240   0.01646  -0.0951   0.1326   1.0000
  12.500   1.7993   0.02418   0.01815  -0.0924   0.1160   1.0000
  12.750   1.8011   0.02593   0.01985  -0.0900   0.1023   1.0000
  13.000   1.7988   0.02809   0.02191  -0.0875   0.0845   1.0000
  13.250   1.7836   0.03137   0.02500  -0.0841   0.0568   1.0000
  13.500   1.7710   0.03462   0.02816  -0.0812   0.0396   1.0000
  13.750   1.7644   0.03745   0.03098  -0.0792   0.0299   1.0000
  14.000   1.7622   0.04001   0.03356  -0.0776   0.0250   1.0000
  14.250   1.7613   0.04254   0.03613  -0.0763   0.0221   1.0000
  14.500   1.7614   0.04502   0.03867  -0.0751   0.0205   1.0000
  14.750   1.7587   0.04787   0.04159  -0.0740   0.0189   1.0000
  15.000   1.7599   0.05041   0.04420  -0.0732   0.0182   1.0000
  15.250   1.7595   0.05317   0.04705  -0.0725   0.0174   1.0000
  15.500   1.7569   0.05625   0.05021  -0.0719   0.0168   1.0000
  15.750   1.7519   0.05971   0.05375  -0.0714   0.0162   1.0000
  16.000   1.7435   0.06372   0.05786  -0.0710   0.0156   1.0000
  16.250   1.7371   0.06758   0.06183  -0.0709   0.0152   1.0000
  16.500   1.7322   0.07128   0.06563  -0.0709   0.0150   1.0000
  16.750   1.7256   0.07529   0.06975  -0.0710   0.0147   1.0000
  17.000   1.7174   0.07962   0.07418  -0.0713   0.0144   1.0000
  17.250   1.7076   0.08425   0.07892  -0.0718   0.0141   1.0000
  17.500   1.6958   0.08920   0.08398  -0.0724   0.0139   1.0000
  17.750   1.6824   0.09448   0.08938  -0.0733   0.0137   1.0000
  18.000   1.6679   0.10003   0.09504  -0.0744   0.0135   1.0000
  18.250   1.6521   0.10590   0.10103  -0.0757   0.0133   1.0000
  18.500   1.6349   0.11204   0.10729  -0.0773   0.0131   1.0000
  18.750   1.6177   0.11830   0.11366  -0.0791   0.0129   1.0000
  19.000   1.5995   0.12484   0.12031  -0.0812   0.0128   1.0000
<< Back to GOE 331 (PFALZ 60) AIRFOIL (goe331-il)

Polar data table (+)

Polar graphs


<< Back to GOE 331 (PFALZ 60) AIRFOIL (goe331-il)