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

GOE 239 (MVA H.31) AIRFOIL (goe239-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 239 (MVA H.31) AIRFOIL (goe239-il)
Reynolds number: 50,000
Max Cl/Cd: 17.32 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe239-il-50000.txt
Download as CSV file: xf-goe239-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 239 (MVA H.31) 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.2594   0.12217   0.11520  -0.0338   1.0000   0.1817
  -9.250  -0.2705   0.12217   0.11533  -0.0334   1.0000   0.1869
  -9.000  -0.2965   0.12433   0.11767  -0.0324   1.0000   0.1883
  -8.750  -0.2640   0.11621   0.10950  -0.0307   1.0000   0.1949
  -8.500  -0.2698   0.11500   0.10839  -0.0291   1.0000   0.2004
  -8.250  -0.2940   0.11637   0.10992  -0.0271   1.0000   0.2035
  -8.000  -0.3263   0.11850   0.11223  -0.0239   1.0000   0.2044
  -7.750  -0.2852   0.10979   0.10344  -0.0233   1.0000   0.2118
  -7.500  -0.2981   0.10923   0.10300  -0.0208   1.0000   0.2173
  -7.250  -0.3276   0.11057   0.10450  -0.0175   1.0000   0.2198
  -7.000  -0.3594   0.11243   0.10655  -0.0168   1.0000   0.2209
  -6.750  -0.3230   0.10469   0.09874  -0.0132   1.0000   0.2309
  -6.500  -0.3455   0.10507   0.09927  -0.0114   1.0000   0.2358
  -6.250  -0.3584   0.10376   0.09808  -0.0110   1.0000   0.2388
  -6.000  -0.3479   0.10006   0.09440  -0.0074   1.0000   0.2460
  -5.750  -0.3649   0.10023   0.09470  -0.0087   1.0000   0.2532
  -5.500  -0.3652   0.09715   0.09170  -0.0069   1.0000   0.2569
  -5.250  -0.3639   0.09483   0.08943  -0.0044   1.0000   0.2634
  -5.000  -0.3735   0.09453   0.08924  -0.0093   1.0000   0.2716
  -4.750  -0.3717   0.09130   0.08608  -0.0045   1.0000   0.2757
  -4.500  -0.3716   0.09115   0.08597  -0.0112   1.0000   0.2877
  -4.250  -0.3730   0.08781   0.08271  -0.0054   1.0000   0.2914
  -4.000  -0.3480   0.08493   0.07980  -0.0114   0.9941   0.3069
  -3.750  -0.3075   0.08106   0.07584  -0.0188   0.9826   0.3240
  -3.500  -0.2665   0.07740   0.07209  -0.0267   0.9710   0.3404
  -3.250  -0.2304   0.07403   0.06865  -0.0327   0.9592   0.3572
  -3.000  -0.1998   0.07107   0.06563  -0.0364   0.9475   0.3771
  -2.750   0.0123   0.05039   0.04300  -0.1024   0.9426   0.1855
  -2.500   0.0725   0.04622   0.03820  -0.1121   0.9328   0.1853
  -2.250   0.1366   0.04264   0.03378  -0.1213   0.9247   0.1894
  -2.000   0.1851   0.04151   0.03237  -0.1260   0.9149   0.2027
  -1.750   0.2219   0.04078   0.03149  -0.1286   0.9038   0.2158
  -1.500   0.2712   0.04006   0.03058  -0.1328   0.8955   0.2427
  -1.250   0.3032   0.04002   0.03049  -0.1342   0.8844   0.2752
  -1.000   0.3352   0.04011   0.03064  -0.1352   0.8743   0.3258
  -0.750   0.3765   0.03959   0.03019  -0.1374   0.8661   0.4198
  -0.500   0.3986   0.03976   0.03039  -0.1369   0.8557   0.4781
  -0.250   0.4415   0.03945   0.03008  -0.1391   0.8483   0.5369
   0.000   0.4562   0.04011   0.03079  -0.1377   0.8372   0.5747
   0.250   0.4893   0.04015   0.03089  -0.1385   0.8289   0.6256
   0.500   0.5158   0.04047   0.03123  -0.1390   0.8191   0.6627
   0.750   0.5406   0.04070   0.03162  -0.1391   0.8097   0.7057
   1.000   0.5635   0.04013   0.03160  -0.1376   0.8010   1.0000
   1.250   0.5880   0.04149   0.03261  -0.1391   0.7898   1.0000
   1.500   0.6303   0.04208   0.03279  -0.1419   0.7812   1.0000
   1.750   0.6528   0.04320   0.03368  -0.1419   0.7704   1.0000
   2.000   0.6712   0.04451   0.03481  -0.1414   0.7597   1.0000
   2.250   0.7182   0.04447   0.03456  -0.1436   0.7517   1.0000
   2.500   0.7254   0.04626   0.03627  -0.1419   0.7394   1.0000
   2.750   0.7437   0.04756   0.03749  -0.1412   0.7284   1.0000
   3.000   0.7908   0.04725   0.03707  -0.1428   0.7202   1.0000
   3.250   0.7970   0.04923   0.03901  -0.1411   0.7074   1.0000
   3.500   0.8109   0.05083   0.04058  -0.1400   0.6957   1.0000
   3.750   0.8641   0.04989   0.03959  -0.1416   0.6883   1.0000
   4.000   0.8669   0.05225   0.04194  -0.1397   0.6749   1.0000
   4.250   0.8729   0.05452   0.04422  -0.1381   0.6625   1.0000
   4.500   0.8958   0.05565   0.04534  -0.1376   0.6520   1.0000
   4.750   0.9284   0.05599   0.04570  -0.1375   0.6424   1.0000
   5.000   0.9253   0.05918   0.04892  -0.1358   0.6295   1.0000
   5.250   0.9356   0.06144   0.05120  -0.1347   0.6186   1.0000
   5.500   0.9712   0.06163   0.05143  -0.1347   0.6101   1.0000
   5.750   0.9546   0.06630   0.05613  -0.1329   0.5978   1.0000
   6.000   0.9690   0.06843   0.05832  -0.1321   0.5885   1.0000
   6.250   0.9795   0.07091   0.06084  -0.1314   0.5790   1.0000
   6.500   0.9667   0.07568   0.06565  -0.1304   0.5706   1.0000
   6.750   0.9742   0.07871   0.06874  -0.1300   0.5633   1.0000
   7.000   0.9758   0.08234   0.07241  -0.1294   0.5565   1.0000
   7.250   0.9591   0.08755   0.07767  -0.1290   0.5511   1.0000
   7.500   0.9569   0.09184   0.08202  -0.1290   0.5485   1.0000
   7.750   0.9402   0.09762   0.08786  -0.1295   0.5513   1.0000
   8.000   0.9351   0.10261   0.09292  -0.1302   0.5554   1.0000
   8.250   0.9507   0.10670   0.09712  -0.1315   0.5593   1.0000
   8.500   0.8571   0.12048   0.11098  -0.1368   0.6554   1.0000
   8.750   0.8633   0.12259   0.11315  -0.1360   0.6429   1.0000
<< Back to GOE 239 (MVA H.31) AIRFOIL (goe239-il)

Polar data table (+)

Polar graphs


<< Back to GOE 239 (MVA H.31) AIRFOIL (goe239-il)