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

GOE 503 AIRFOIL (goe503-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 503 AIRFOIL (goe503-il)
Reynolds number: 100,000
Max Cl/Cd: 32.3 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe503-il-100000.txt
Download as CSV file: xf-goe503-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 503 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.1002   0.11230   0.10796  -0.0770   0.9455   0.1109
  -9.250  -0.1090   0.11032   0.10601  -0.0828   0.9351   0.1149
  -9.000  -0.1187   0.10705   0.10280  -0.0923   0.9279   0.1160
  -8.750  -0.0563   0.10031   0.09594  -0.0892   0.9281   0.1220
  -8.500  -0.0542   0.09783   0.09347  -0.0909   0.9169   0.1271
  -8.250  -0.0837   0.09653   0.09224  -0.0966   0.9028   0.1301
  -8.000  -0.0485   0.09042   0.08609  -0.0982   0.9004   0.1330
  -7.750  -0.0166   0.08621   0.08182  -0.1014   0.8977   0.1400
  -7.500  -0.0386   0.08445   0.08010  -0.1029   0.8836   0.1439
  -7.000  -0.0173   0.07750   0.07309  -0.1047   0.8683   0.1547
  -6.750  -0.0478   0.07620   0.07171  -0.1067   0.8531   0.1606
  -6.500  -0.1161   0.05976   0.05475  -0.1080   0.8376   0.1014
  -6.250  -0.0655   0.06068   0.05583  -0.1097   0.8358   0.1183
  -6.000  -0.0603   0.06271   0.05778  -0.1069   0.8244   0.1465
  -5.750  -0.0286   0.05877   0.05381  -0.1083   0.8212   0.1494
  -5.500  -0.0579   0.05146   0.04605  -0.1046   0.8086   0.1242
  -5.250  -0.0490   0.04472   0.03852  -0.1050   0.8039   0.1292
  -5.000  -0.0369   0.04568   0.03968  -0.1021   0.7940   0.1356
  -4.750  -0.0110   0.04244   0.03602  -0.1028   0.7894   0.1414
  -4.500   0.0171   0.03764   0.03030  -0.1039   0.7866   0.1463
  -4.250   0.0119   0.03760   0.03028  -0.0985   0.7747   0.1479
  -4.000   0.0584   0.03618   0.02873  -0.1016   0.7721   0.1553
  -3.750   0.0589   0.03546   0.02770  -0.0971   0.7622   0.1582
  -3.500   0.0929   0.03345   0.02525  -0.0980   0.7577   0.1637
  -3.250   0.1405   0.03221   0.02386  -0.1011   0.7552   0.1731
  -3.000   0.1356   0.03239   0.02371  -0.0955   0.7443   0.1774
  -2.750   0.1794   0.03123   0.02256  -0.0981   0.7408   0.1863
  -2.500   0.2271   0.03000   0.02112  -0.1012   0.7382   0.2006
  -2.250   0.2186   0.03079   0.02186  -0.0951   0.7269   0.2057
  -2.000   0.2624   0.02980   0.02073  -0.0975   0.7236   0.2206
  -1.750   0.3117   0.02881   0.01959  -0.1009   0.7210   0.2389
  -1.500   0.2961   0.02994   0.02064  -0.0937   0.7093   0.2440
  -1.250   0.3417   0.02909   0.01971  -0.0965   0.7062   0.2628
  -1.000   0.3388   0.02988   0.02059  -0.0915   0.6970   0.2702
  -0.750   0.3691   0.02957   0.02023  -0.0918   0.6917   0.2865
  -0.500   0.4179   0.02870   0.01928  -0.0950   0.6889   0.3086
  -0.250   0.3937   0.03029   0.02086  -0.0867   0.6777   0.3124
   0.000   0.4398   0.02947   0.02005  -0.0896   0.6743   0.3331
   0.250   0.4927   0.02851   0.01907  -0.0935   0.6717   0.3595
   0.500   0.4555   0.03053   0.02116  -0.0832   0.6598   0.3598
   0.750   0.5144   0.02945   0.02014  -0.0882   0.6572   0.3898
   1.000   0.5780   0.02826   0.01900  -0.0941   0.6549   0.4305
   1.250   0.8076   0.02578   0.01775  -0.1326   0.6512   1.0000
   1.500   0.8020   0.02675   0.01869  -0.1270   0.6433   1.0000
   1.750   0.8397   0.02654   0.01827  -0.1285   0.6391   1.0000
   2.000   0.8242   0.02785   0.01961  -0.1212   0.6313   1.0000
   2.250   0.8430   0.02814   0.01979  -0.1196   0.6256   1.0000
   2.500   0.8879   0.02772   0.01920  -0.1223   0.6221   1.0000
   2.750   0.8465   0.02981   0.02139  -0.1109   0.6126   1.0000
   3.000   0.8815   0.02964   0.02110  -0.1120   0.6084   1.0000
   3.250   0.9338   0.02902   0.02032  -0.1158   0.6054   1.0000
   3.500   0.8700   0.03188   0.02332  -0.1011   0.5945   1.0000
   3.750   0.9207   0.03117   0.02249  -0.1046   0.5914   1.0000
   4.000   0.9802   0.03035   0.02152  -0.1095   0.5888   1.0000
   4.250   0.8856   0.03421   0.02554  -0.0903   0.5767   1.0000
   4.500   0.9591   0.03284   0.02405  -0.0972   0.5746   1.0000
   4.750   0.8483   0.03822   0.02956  -0.0776   0.5603   1.0000
   5.000   0.9142   0.03643   0.02767  -0.0821   0.5593   1.0000
   5.250   0.9885   0.03481   0.02596  -0.0887   0.5578   1.0000
   5.500   0.8363   0.04405   0.03537  -0.0672   0.5380   1.0000
   5.750   0.9061   0.04116   0.03241  -0.0707   0.5387   1.0000
   6.000   0.9975   0.03786   0.02902  -0.0779   0.5396   1.0000
   6.250   0.9140   0.04406   0.03531  -0.0656   0.5254   1.0000
   6.500   0.9805   0.04151   0.03271  -0.0690   0.5249   1.0000
   6.750   1.0617   0.03879   0.02993  -0.0750   0.5242   1.0000
   7.000   0.7438   0.06644   0.05786  -0.0513   0.4788   1.0000
   7.250   0.7183   0.07154   0.06299  -0.0493   0.4680   1.0000
   7.500   0.7432   0.07168   0.06312  -0.0487   0.4636   1.0000
   7.750   0.7849   0.07008   0.06148  -0.0485   0.4615   1.0000
   8.000   0.7490   0.07635   0.06781  -0.0466   0.4487   1.0000
   8.250   0.7882   0.07488   0.06632  -0.0462   0.4461   1.0000
   8.500   0.7597   0.08055   0.07203  -0.0448   0.4342   1.0000
   8.750   0.7936   0.07959   0.07106  -0.0442   0.4309   1.0000
   9.000   0.8121   0.08031   0.07179  -0.0434   0.4258   1.0000
   9.250   0.8814   0.07522   0.06669  -0.0432   0.4280   1.0000
   9.500   0.8379   0.08286   0.07439  -0.0419   0.4138   1.0000
   9.750   0.8068   0.08917   0.08075  -0.0411   0.4011   1.0000
  10.000   0.8414   0.08801   0.07961  -0.0404   0.3986   1.0000
  10.250   0.8819   0.08606   0.07768  -0.0397   0.3972   1.0000
  10.500   0.8445   0.09338   0.08505  -0.0392   0.3836   1.0000
  10.750   0.8857   0.09116   0.08286  -0.0384   0.3819   1.0000
  11.000   0.8501   0.09852   0.09027  -0.0382   0.3686   1.0000
  11.250   0.8860   0.09699   0.08877  -0.0374   0.3668   1.0000
  11.500   0.8451   0.10552   0.09735  -0.0379   0.3539   1.0000
  11.750   0.8778   0.10440   0.09628  -0.0370   0.3517   1.0000
  12.000   0.9132   0.10287   0.09478  -0.0361   0.3503   1.0000
  12.250   0.8680   0.11224   0.10421  -0.0371   0.3370   1.0000
  12.500   0.9036   0.11060   0.10261  -0.0361   0.3354   1.0000
  12.750   0.5825   0.14072   0.13369  -0.0344   0.3658   1.0000
  13.000   0.5332   0.14013   0.13320  -0.0349   0.3584   1.0000
<< Back to GOE 503 AIRFOIL (goe503-il)

Polar data table (+)

Polar graphs


<< Back to GOE 503 AIRFOIL (goe503-il)