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

GOE 190 (MVA MK.18) AIRFOIL (goe190-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 190 (MVA MK.18) AIRFOIL (goe190-il)
Reynolds number: 50,000
Max Cl/Cd: 37.2 at α=6.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe190-il-50000-n5.txt
Download as CSV file: xf-goe190-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 190 (MVA MK.18) 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3179   0.10645   0.09940  -0.0376   1.0000   0.0797
  -8.750  -0.3224   0.10345   0.09649  -0.0374   1.0000   0.0797
  -8.500  -0.3296   0.10043   0.09355  -0.0372   1.0000   0.0801
  -8.250  -0.3380   0.09748   0.09070  -0.0367   1.0000   0.0802
  -8.000  -0.3486   0.09468   0.08801  -0.0359   1.0000   0.0803
  -7.750  -0.3602   0.09212   0.08556  -0.0346   1.0000   0.0800
  -7.500  -0.3726   0.08934   0.08290  -0.0335   1.0000   0.0798
  -7.250  -0.3828   0.08616   0.07982  -0.0333   1.0000   0.0794
  -7.000  -0.3932   0.08257   0.07633  -0.0339   1.0000   0.0790
  -6.750  -0.4035   0.07822   0.07207  -0.0356   1.0000   0.0788
  -6.500  -0.3902   0.06611   0.05982  -0.0524   0.9914   0.0796
  -6.250  -0.3426   0.04635   0.03882  -0.0868   0.9831   0.0813
  -6.000  -0.3014   0.04117   0.03290  -0.0955   0.9758   0.0828
  -5.750  -0.2670   0.03921   0.03085  -0.0985   0.9684   0.0861
  -5.500  -0.2293   0.03709   0.02842  -0.1024   0.9607   0.0901
  -5.250  -0.1934   0.03513   0.02604  -0.1053   0.9522   0.0930
  -5.000  -0.1544   0.03354   0.02408  -0.1083   0.9447   0.0963
  -4.750  -0.1247   0.03253   0.02315  -0.1093   0.9344   0.1010
  -4.500  -0.0882   0.03149   0.02192  -0.1116   0.9260   0.1069
  -4.250  -0.0554   0.03053   0.02088  -0.1129   0.9155   0.1120
  -4.000  -0.0236   0.02967   0.02000  -0.1142   0.9044   0.1202
  -3.750   0.0123   0.02872   0.01908  -0.1164   0.8947   0.1340
  -3.500   0.0485   0.02781   0.01835  -0.1184   0.8848   0.1570
  -3.250   0.0775   0.02809   0.01927  -0.1181   0.8724   0.2097
  -3.000   0.1104   0.02870   0.01969  -0.1183   0.8603   0.2698
  -2.750   0.1438   0.02941   0.02021  -0.1184   0.8492   0.3033
  -2.500   0.1786   0.03020   0.02092  -0.1183   0.8401   0.3306
  -2.250   0.2069   0.03075   0.02141  -0.1174   0.8279   0.3508
  -2.000   0.2382   0.03117   0.02174  -0.1170   0.8175   0.3739
  -1.750   0.2759   0.03106   0.02151  -0.1179   0.8094   0.3916
  -1.500   0.3086   0.03066   0.02096  -0.1186   0.7978   0.3995
  -1.250   0.3508   0.03000   0.02006  -0.1211   0.7896   0.4073
  -1.000   0.3863   0.02952   0.01937  -0.1226   0.7781   0.4136
  -0.750   0.4205   0.02913   0.01884  -0.1236   0.7670   0.4191
  -0.500   0.4621   0.02858   0.01803  -0.1260   0.7578   0.4258
  -0.250   0.4930   0.02835   0.01767  -0.1265   0.7450   0.4307
   0.000   0.5267   0.02809   0.01728  -0.1274   0.7338   0.4360
   0.250   0.5636   0.02780   0.01678  -0.1289   0.7232   0.4419
   0.500   0.5918   0.02775   0.01661  -0.1291   0.7104   0.4463
   0.750   0.6223   0.02766   0.01644  -0.1294   0.6991   0.4512
   1.000   0.6560   0.02757   0.01619  -0.1303   0.6886   0.4573
   1.250   0.6821   0.02770   0.01623  -0.1302   0.6764   0.4620
   1.500   0.7115   0.02773   0.01619  -0.1304   0.6659   0.4664
   1.750   0.7420   0.02778   0.01613  -0.1308   0.6558   0.4723
   2.000   0.7673   0.02800   0.01629  -0.1305   0.6447   0.4778
   2.250   0.7980   0.02803   0.01626  -0.1308   0.6357   0.4834
   2.500   0.8241   0.02825   0.01644  -0.1306   0.6252   0.4890
   2.750   0.8510   0.02845   0.01661  -0.1304   0.6155   0.4945
   3.000   0.8812   0.02853   0.01664  -0.1307   0.6068   0.5010
   3.250   0.9059   0.02887   0.01699  -0.1303   0.5968   0.5078
   3.500   0.9388   0.02886   0.01694  -0.1309   0.5891   0.5146
   3.750   0.9614   0.02931   0.01743  -0.1302   0.5789   0.5215
   4.000   0.9921   0.02941   0.01754  -0.1305   0.5711   0.5288
   4.250   1.0159   0.02985   0.01802  -0.1299   0.5615   0.5371
   4.500   1.0448   0.03005   0.01826  -0.1300   0.5535   0.5456
   4.750   1.0689   0.03051   0.01881  -0.1294   0.5443   0.5546
   5.000   1.0968   0.03080   0.01916  -0.1293   0.5363   0.5644
   5.250   1.1197   0.03129   0.01978  -0.1286   0.5271   0.5748
   5.500   1.1471   0.03164   0.02023  -0.1285   0.5193   0.5876
   5.750   1.1680   0.03222   0.02099  -0.1275   0.5100   0.6009
   6.000   1.1942   0.03257   0.02148  -0.1271   0.5018   0.6181
   6.250   1.2134   0.03307   0.02228  -0.1259   0.4919   0.6404
   6.750   1.2530   0.03368   0.02337  -0.1229   0.4726   1.0000
   7.000   1.2683   0.03456   0.02433  -0.1211   0.4623   1.0000
   7.250   1.2958   0.03490   0.02464  -0.1207   0.4536   1.0000
   7.500   1.3046   0.03608   0.02598  -0.1182   0.4434   1.0000
   7.750   1.3312   0.03643   0.02632  -0.1176   0.4354   1.0000
   8.000   1.3381   0.03767   0.02773  -0.1149   0.4254   1.0000
   8.250   1.3581   0.03828   0.02842  -0.1135   0.4171   1.0000
   8.500   1.3691   0.03928   0.02956  -0.1112   0.4078   1.0000
   8.750   1.3816   0.04021   0.03059  -0.1091   0.3989   1.0000
   9.000   1.3986   0.04084   0.03131  -0.1074   0.3901   1.0000
   9.250   1.4018   0.04217   0.03281  -0.1042   0.3809   1.0000
   9.500   1.4272   0.04228   0.03293  -0.1033   0.3723   1.0000
   9.750   1.4186   0.04417   0.03502  -0.0990   0.3627   1.0000
  10.000   1.4350   0.04465   0.03554  -0.0972   0.3534   1.0000
  10.250   1.4375   0.04587   0.03687  -0.0941   0.3431   1.0000
  10.500   1.4350   0.04750   0.03862  -0.0909   0.3327   1.0000
  10.750   1.4502   0.04788   0.03902  -0.0890   0.3220   1.0000
  11.000   1.4413   0.05016   0.04146  -0.0858   0.3118   1.0000
  11.250   1.4391   0.05213   0.04352  -0.0834   0.3018   1.0000
  11.500   1.4553   0.05255   0.04395  -0.0818   0.2916   1.0000
  11.750   1.4296   0.05701   0.04863  -0.0794   0.2825   1.0000
  12.000   1.4369   0.05842   0.05009  -0.0778   0.2731   1.0000
  12.250   1.4214   0.06244   0.05428  -0.0766   0.2639   1.0000
  12.500   1.4131   0.06590   0.05785  -0.0757   0.2549   1.0000
  12.750   1.4210   0.06732   0.05933  -0.0745   0.2453   1.0000
  13.000   1.3888   0.07444   0.06666  -0.0752   0.2370   1.0000
  13.250   1.4003   0.07545   0.06769  -0.0741   0.2279   1.0000
  13.500   1.3713   0.08274   0.07517  -0.0755   0.2195   1.0000
  13.750   1.3675   0.08630   0.07880  -0.0756   0.2109   1.0000
  14.000   1.3696   0.08888   0.08144  -0.0753   0.2021   1.0000
  14.250   1.3393   0.09740   0.09011  -0.0780   0.1944   1.0000
  14.500   1.3744   0.09406   0.08669  -0.0749   0.1852   1.0000
  14.750   1.3151   0.10851   0.10140  -0.0810   0.1785   1.0000
  15.000   1.3583   0.10335   0.09615  -0.0769   0.1700   1.0000
  15.250   1.2823   0.12222   0.11526  -0.0861   0.1637   1.0000
  15.500   1.3452   0.11243   0.10539  -0.0793   0.1553   1.0000
  15.750   1.2390   0.13966   0.13280  -0.0941   0.1492   1.0000
  16.000   1.2822   0.13306   0.12629  -0.0891   0.1432   1.0000
<< Back to GOE 190 (MVA MK.18) AIRFOIL (goe190-il)

Polar data table (+)

Polar graphs


<< Back to GOE 190 (MVA MK.18) AIRFOIL (goe190-il)