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

GOE 491 AIRFOIL (goe491-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 491 AIRFOIL (goe491-il)
Reynolds number: 50,000
Max Cl/Cd: 34.63 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe491-il-50000.txt
Download as CSV file: xf-goe491-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 491 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4980   0.11217   0.10510   0.0022   1.0000   0.1221
  -8.500  -0.5028   0.11059   0.10362  -0.0002   1.0000   0.1251
  -8.250  -0.5138   0.10994   0.10310  -0.0035   1.0000   0.1262
  -8.000  -0.4904   0.10265   0.09578  -0.0006   1.0000   0.1317
  -7.750  -0.4904   0.09994   0.09315  -0.0017   1.0000   0.1369
  -7.500  -0.4977   0.09882   0.09212  -0.0058   1.0000   0.1398
  -7.250  -0.4865   0.09355   0.08690  -0.0042   1.0000   0.1442
  -7.000  -0.4813   0.09047   0.08382  -0.0053   1.0000   0.1513
  -6.750  -0.4808   0.08805   0.08145  -0.0095   1.0000   0.1556
  -6.500  -0.4695   0.08366   0.07709  -0.0074   1.0000   0.1645
  -6.000  -0.4576   0.07847   0.07187  -0.0132   1.0000   0.1830
  -5.750  -0.4476   0.07436   0.06781  -0.0119   1.0000   0.1958
  -5.500  -0.4389   0.07132   0.06476  -0.0130   1.0000   0.2110
  -5.250  -0.4300   0.06686   0.06041  -0.0098   1.0000   0.2290
  -5.000  -0.4223   0.06378   0.05734  -0.0093   1.0000   0.2550
  -4.750  -0.4146   0.06047   0.05409  -0.0074   1.0000   0.2839
  -4.500  -0.4089   0.05765   0.05131  -0.0053   1.0000   0.3239
  -4.000  -0.4055   0.05089   0.04484   0.0066   1.0000   0.4270
  -3.750  -0.4057   0.04777   0.04188   0.0138   1.0000   0.4847
  -3.500  -0.4042   0.04500   0.03921   0.0189   1.0000   0.5340
  -3.250  -0.4016   0.04192   0.03630   0.0272   1.0000   0.5858
  -3.000  -0.3956   0.03919   0.03362   0.0303   1.0000   0.6189
  -2.500  -0.1951   0.03275   0.02360  -0.0167   1.0000   0.1816
  -2.250  -0.1707   0.03029   0.02077  -0.0157   1.0000   0.1806
  -2.000  -0.1451   0.02791   0.01797  -0.0146   1.0000   0.1768
  -1.750  -0.1201   0.02603   0.01567  -0.0135   1.0000   0.1853
  -1.500  -0.0941   0.02430   0.01363  -0.0125   1.0000   0.1922
  -1.250  -0.0686   0.02293   0.01201  -0.0117   1.0000   0.2071
  -1.000  -0.0398   0.02167   0.01049  -0.0112   1.0000   0.2146
  -0.750  -0.0066   0.02072   0.00925  -0.0115   1.0000   0.2253
  -0.500   0.0230   0.01978   0.00832  -0.0116   1.0000   0.2466
  -0.250   0.0493   0.01904   0.00757  -0.0110   1.0000   0.2634
   0.000   0.0723   0.01833   0.00701  -0.0100   1.0000   0.2885
   0.250   0.0942   0.01702   0.00653  -0.0092   1.0000   0.3971
   0.500   0.1619   0.01583   0.00579  -0.0157   1.0000   1.0000
   0.750   0.1795   0.01608   0.00586  -0.0142   1.0000   1.0000
   1.000   0.1961   0.01638   0.00604  -0.0127   1.0000   1.0000
   1.250   0.2115   0.01677   0.00636  -0.0112   1.0000   1.0000
   1.500   0.2251   0.01728   0.00685  -0.0098   1.0000   1.0000
   1.750   0.2360   0.01800   0.00756  -0.0086   1.0000   1.0000
   2.000   0.2582   0.01909   0.00868  -0.0108   0.9933   1.0000
   2.250   0.3644   0.01993   0.00969  -0.0277   0.9452   1.0000
   2.500   0.4662   0.01980   0.00986  -0.0418   0.9004   1.0000
   2.750   0.5373   0.01948   0.00975  -0.0486   0.8540   1.0000
   3.000   0.5819   0.01935   0.00977  -0.0499   0.8093   1.0000
   3.250   0.6130   0.01945   0.00989  -0.0487   0.7689   1.0000
   3.500   0.6411   0.01968   0.01010  -0.0471   0.7330   1.0000
   3.750   0.6647   0.02016   0.01061  -0.0452   0.6979   1.0000
   4.000   0.6889   0.02068   0.01115  -0.0434   0.6655   1.0000
   4.250   0.7123   0.02129   0.01184  -0.0414   0.6341   1.0000
   4.500   0.7325   0.02188   0.01247  -0.0389   0.5974   1.0000
   4.750   0.7518   0.02231   0.01285  -0.0359   0.5571   1.0000
   5.000   0.7713   0.02279   0.01329  -0.0331   0.5178   1.0000
   5.250   0.7878   0.02303   0.01341  -0.0296   0.4699   1.0000
   5.500   0.8015   0.02318   0.01344  -0.0258   0.4139   1.0000
   5.750   0.8148   0.02353   0.01372  -0.0223   0.3519   1.0000
   6.000   0.8197   0.02476   0.01424  -0.0173   0.2304   1.0000
   6.250   0.8384   0.02731   0.01628  -0.0151   0.1768   1.0000
   6.500   0.8594   0.02938   0.01826  -0.0136   0.1507   1.0000
   6.750   0.8835   0.03181   0.02079  -0.0123   0.1369   1.0000
   7.000   0.9063   0.03449   0.02377  -0.0110   0.1271   1.0000
   7.250   0.9263   0.03702   0.02643  -0.0097   0.1164   1.0000
   7.500   0.9444   0.04040   0.03043  -0.0079   0.1123   1.0000
   7.750   0.9600   0.04410   0.03462  -0.0061   0.1097   1.0000
   8.000   0.9758   0.04777   0.03836  -0.0049   0.1048   1.0000
   8.250   0.9803   0.05193   0.04324  -0.0028   0.1031   1.0000
   8.500   0.9825   0.05664   0.04845  -0.0011   0.1030   1.0000
   8.750   0.9843   0.06165   0.05381   0.0001   0.1041   1.0000
   9.000   0.9842   0.06686   0.05934   0.0010   0.1058   1.0000
   9.250   0.9393   0.07398   0.06704   0.0002   0.1111   1.0000
   9.500   0.9174   0.07993   0.07308  -0.0009   0.1141   1.0000
   9.750   0.9067   0.08536   0.07852  -0.0017   0.1167   1.0000
  10.000   0.8662   0.09696   0.09004  -0.0117   0.1298   1.0000
  10.250   0.8699   0.10231   0.09538  -0.0121   0.1314   1.0000
<< Back to GOE 491 AIRFOIL (goe491-il)

Polar data table (+)

Polar graphs


<< Back to GOE 491 AIRFOIL (goe491-il)