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

GOE 241 (MVA PR.1) AIRFOIL (goe241-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 241 (MVA PR.1) AIRFOIL (goe241-il)
Reynolds number: 200,000
Max Cl/Cd: 67.09 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe241-il-200000-n5.txt
Download as CSV file: xf-goe241-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 241 (MVA PR.1) 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000   0.1130   0.08876   0.08422  -0.1157   0.8035   0.0404
  -9.750   0.1233   0.08642   0.08184  -0.1164   0.7969   0.0402
  -9.500   0.1249   0.08218   0.07756  -0.1189   0.7914   0.0403
  -9.250   0.1263   0.07782   0.07320  -0.1217   0.7849   0.0404
  -9.000   0.1318   0.07449   0.06985  -0.1235   0.7785   0.0402
  -8.500   0.1383   0.06677   0.06210  -0.1282   0.7658   0.0398
  -8.250   0.1256   0.05890   0.05420  -0.1340   0.7594   0.0394
  -8.000   0.0865   0.04000   0.03473  -0.1648   0.7531   0.0388
  -7.750   0.1229   0.03299   0.02703  -0.1831   0.7450   0.0397
  -7.500   0.1574   0.02968   0.02318  -0.1901   0.7379   0.0402
  -7.250   0.1908   0.02745   0.02053  -0.1942   0.7312   0.0406
  -7.000   0.2198   0.02617   0.01915  -0.1958   0.7233   0.0409
  -6.750   0.2501   0.02508   0.01790  -0.1973   0.7165   0.0413
  -6.500   0.2798   0.02414   0.01685  -0.1985   0.7088   0.0417
  -6.250   0.3100   0.02324   0.01580  -0.1997   0.7007   0.0422
  -6.000   0.3404   0.02242   0.01482  -0.2008   0.6928   0.0429
  -5.750   0.3705   0.02169   0.01394  -0.2017   0.6835   0.0438
  -5.500   0.4010   0.02100   0.01304  -0.2026   0.6748   0.0449
  -5.250   0.4310   0.02037   0.01222  -0.2033   0.6649   0.0456
  -5.000   0.4602   0.01977   0.01152  -0.2037   0.6560   0.0462
  -4.750   0.4889   0.01928   0.01100  -0.2041   0.6461   0.0468
  -4.500   0.5178   0.01888   0.01054  -0.2044   0.6369   0.0475
  -4.250   0.5466   0.01852   0.01012  -0.2048   0.6269   0.0484
  -4.000   0.5756   0.01821   0.00974  -0.2051   0.6175   0.0495
  -3.750   0.6046   0.01797   0.00939  -0.2054   0.6075   0.0509
  -3.500   0.6337   0.01772   0.00908  -0.2058   0.5979   0.0523
  -3.250   0.6628   0.01750   0.00882  -0.2063   0.5882   0.0538
  -3.000   0.6922   0.01731   0.00860  -0.2068   0.5790   0.0552
  -2.750   0.7217   0.01718   0.00838  -0.2073   0.5701   0.0569
  -2.500   0.7514   0.01708   0.00819  -0.2079   0.5623   0.0586
  -2.250   0.7818   0.01693   0.00800  -0.2087   0.5541   0.0613
  -2.000   0.8110   0.01690   0.00785  -0.2092   0.5464   0.0648
  -1.750   0.8405   0.01682   0.00770  -0.2098   0.5379   0.0684
  -1.500   0.8689   0.01684   0.00758  -0.2101   0.5301   0.0729
  -1.250   0.8980   0.01680   0.00750  -0.2105   0.5226   0.0808
  -1.000   0.9278   0.01667   0.00741  -0.2113   0.5156   0.1122
  -0.750   0.9602   0.01627   0.00756  -0.2131   0.5098   0.3359
  -0.500   0.9873   0.01644   0.00776  -0.2130   0.5042   0.3693
  -0.250   1.0137   0.01666   0.00794  -0.2127   0.4988   0.3909
   0.000   1.0395   0.01693   0.00816  -0.2123   0.4939   0.4056
   0.250   1.0655   0.01717   0.00838  -0.2119   0.4891   0.4185
   0.500   1.0915   0.01739   0.00855  -0.2116   0.4839   0.4309
   0.750   1.1165   0.01766   0.00883  -0.2111   0.4792   0.4407
   1.000   1.1420   0.01795   0.00901  -0.2107   0.4750   0.4518
   1.250   1.1671   0.01820   0.00927  -0.2102   0.4710   0.4583
   1.500   1.1922   0.01842   0.00949  -0.2097   0.4665   0.4639
   1.750   1.2168   0.01865   0.00967  -0.2093   0.4615   0.4697
   2.000   1.2407   0.01892   0.00986  -0.2087   0.4569   0.4751
   2.250   1.2646   0.01919   0.01012  -0.2080   0.4531   0.4796
   2.500   1.2889   0.01943   0.01039  -0.2075   0.4496   0.4850
   2.750   1.3127   0.01968   0.01063  -0.2069   0.4455   0.4907
   3.000   1.3353   0.01996   0.01087  -0.2061   0.4411   0.4956
   3.250   1.3568   0.02028   0.01116  -0.2051   0.4368   0.4997
   3.500   1.3784   0.02057   0.01147  -0.2042   0.4328   0.5048
   3.750   1.3988   0.02085   0.01177  -0.2030   0.4288   0.5107
   4.000   1.4183   0.02116   0.01208  -0.2017   0.4250   0.5158
   4.250   1.4377   0.02151   0.01244  -0.2004   0.4213   0.5196
   4.500   1.4577   0.02190   0.01280  -0.1993   0.4179   0.5242
   4.750   1.4783   0.02227   0.01320  -0.1983   0.4143   0.5297
   5.000   1.4983   0.02266   0.01363  -0.1973   0.4103   0.5351
   5.250   1.5171   0.02308   0.01410  -0.1961   0.4062   0.5390
   5.500   1.5352   0.02354   0.01457  -0.1947   0.4021   0.5436
   5.750   1.5531   0.02405   0.01505  -0.1935   0.3981   0.5487
   6.000   1.5709   0.02454   0.01561  -0.1922   0.3934   0.5537
   6.250   1.5873   0.02508   0.01622  -0.1908   0.3884   0.5577
   6.500   1.6025   0.02570   0.01685  -0.1893   0.3834   0.5622
   6.750   1.6182   0.02634   0.01751  -0.1879   0.3787   0.5671
   7.000   1.6336   0.02700   0.01824  -0.1865   0.3730   0.5718
   7.250   1.6472   0.02775   0.01903  -0.1849   0.3674   0.5758
   7.500   1.6602   0.02858   0.01988  -0.1833   0.3620   0.5802
   7.750   1.6738   0.02941   0.02079  -0.1819   0.3553   0.5850
   8.000   1.6842   0.03044   0.02182  -0.1801   0.3486   0.5893
   8.250   1.6958   0.03145   0.02290  -0.1786   0.3413   0.5933
   8.500   1.7046   0.03266   0.02414  -0.1769   0.3332   0.5978
   8.750   1.7135   0.03392   0.02543  -0.1752   0.3248   0.6022
   9.000   1.7189   0.03547   0.02698  -0.1733   0.3159   0.6062
   9.250   1.7256   0.03697   0.02853  -0.1717   0.3067   0.6101
   9.500   1.7285   0.03880   0.03035  -0.1697   0.2982   0.6144
   9.750   1.7326   0.04063   0.03220  -0.1680   0.2900   0.6190
  10.000   1.7338   0.04275   0.03432  -0.1662   0.2833   0.6231
  10.250   1.7385   0.04463   0.03625  -0.1647   0.2774   0.6270
  10.500   1.7410   0.04676   0.03840  -0.1631   0.2721   0.6314
  10.750   1.7423   0.04904   0.04068  -0.1616   0.2680   0.6362
  11.000   1.7488   0.05091   0.04262  -0.1604   0.2644   0.6414
  11.250   1.7540   0.05291   0.04469  -0.1593   0.2610   0.6468
  11.500   1.7587   0.05498   0.04681  -0.1581   0.2580   0.6526
  11.750   1.7630   0.05711   0.04899  -0.1570   0.2554   0.6584
  12.000   1.7677   0.05919   0.05110  -0.1559   0.2529   0.6653
  12.250   1.7743   0.06114   0.05312  -0.1550   0.2507   0.6742
  12.500   1.7815   0.06307   0.05518  -0.1542   0.2484   0.6860
  12.750   1.7872   0.06516   0.05740  -0.1534   0.2458   0.7036
  13.000   1.7930   0.06718   0.05961  -0.1526   0.2433   0.7562
  13.250   1.7940   0.06915   0.06171  -0.1512   0.2411   1.0000
  13.500   1.7996   0.07127   0.06384  -0.1505   0.2389   1.0000
  13.750   1.8071   0.07312   0.06569  -0.1498   0.2370   1.0000
  14.000   1.8166   0.07471   0.06727  -0.1491   0.2351   1.0000
  14.250   1.8213   0.07707   0.06976  -0.1486   0.2335   1.0000
  14.500   1.8251   0.07953   0.07233  -0.1481   0.2314   1.0000
  14.750   1.8276   0.08217   0.07506  -0.1477   0.2292   1.0000
  15.000   1.8300   0.08481   0.07779  -0.1473   0.2268   1.0000
  15.250   1.8327   0.08739   0.08042  -0.1469   0.2242   1.0000
  15.500   1.8372   0.08965   0.08269  -0.1465   0.2216   1.0000
  15.750   1.8463   0.09122   0.08423  -0.1460   0.2191   1.0000
  16.000   1.8477   0.09405   0.08717  -0.1458   0.2172   1.0000
  16.250   1.8461   0.09733   0.09061  -0.1458   0.2152   1.0000
  16.500   1.8433   0.10079   0.09421  -0.1459   0.2128   1.0000
<< Back to GOE 241 (MVA PR.1) AIRFOIL (goe241-il)

Polar data table (+)

Polar graphs


<< Back to GOE 241 (MVA PR.1) AIRFOIL (goe241-il)