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

EPPLER 694 AIRFOIL (e694-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 694 AIRFOIL (e694-il)
Reynolds number: 200,000
Max Cl/Cd: 61.67 at α=9°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e694-il-200000.txt
Download as CSV file: xf-e694-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 694 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.2073   0.09888   0.09536  -0.1034   0.9102   0.0577
  -8.500  -0.2043   0.09530   0.09179  -0.1058   0.9085   0.0592
  -8.250  -0.2282   0.09406   0.09061  -0.1015   0.9028   0.0599
  -8.000  -0.2485   0.09189   0.08849  -0.0993   0.8982   0.0606
  -7.750  -0.2662   0.08795   0.08459  -0.1008   0.8953   0.0616
  -7.500  -0.2902   0.08239   0.07901  -0.1057   0.8927   0.0623
  -7.250  -0.3336   0.08387   0.08059  -0.0943   0.8861   0.0608
  -7.000  -0.3554   0.07992   0.07658  -0.0949   0.8823   0.0615
  -6.750  -0.3781   0.07476   0.07064  -0.1006   0.8786   0.0651
  -6.500  -0.3859   0.06871   0.06467  -0.0992   0.8757   0.0663
  -6.250  -0.3929   0.06694   0.06296  -0.0953   0.8722   0.0670
  -6.000  -0.3831   0.06474   0.06077  -0.0944   0.8694   0.0683
  -5.750  -0.3689   0.06225   0.05820  -0.0948   0.8675   0.0705
  -5.250  -0.3072   0.04599   0.04014  -0.1029   0.8650   0.0454
  -5.000  -0.2618   0.04092   0.03386  -0.1056   0.8642   0.0347
  -4.750  -0.2226   0.03819   0.03077  -0.1085   0.8634   0.0353
  -4.500  -0.2254   0.03761   0.03014  -0.1040   0.8578   0.0358
  -4.250  -0.2013   0.03669   0.02911  -0.1040   0.8552   0.0367
  -4.000  -0.1702   0.03566   0.02793  -0.1048   0.8530   0.0372
  -3.750  -0.1358   0.03477   0.02692  -0.1061   0.8514   0.0379
  -3.500  -0.0992   0.03403   0.02613  -0.1079   0.8501   0.0395
  -3.250  -0.0610   0.03347   0.02550  -0.1100   0.8491   0.0420
  -3.000  -0.0592   0.03347   0.02549  -0.1063   0.8423   0.0434
  -2.750  -0.0281   0.03304   0.02510  -0.1077   0.8397   0.0510
  -2.500   0.0112   0.03234   0.02449  -0.1106   0.8379   0.0729
  -2.250   0.0620   0.03070   0.02423  -0.1173   0.8371   0.3600
  -2.000   0.0953   0.03005   0.02489  -0.1183   0.8358   0.6553
  -1.750   0.0980   0.03077   0.02574  -0.1134   0.8296   0.7230
  -1.500   0.1153   0.03137   0.02629  -0.1108   0.8254   0.7650
  -1.250   0.1395   0.03175   0.02659  -0.1091   0.8230   0.7867
  -1.000   0.1670   0.03205   0.02679  -0.1080   0.8214   0.8053
  -0.750   0.1631   0.03270   0.02745  -0.1025   0.8129   0.8176
  -0.500   0.1834   0.03286   0.02755  -0.1002   0.8101   0.8336
  -0.250   0.2083   0.03287   0.02747  -0.0986   0.8083   0.8481
   0.000   0.2068   0.03336   0.02795  -0.0936   0.8000   0.8607
   0.250   0.2206   0.03324   0.02780  -0.0898   0.7968   0.8790
   0.500   0.2416   0.03293   0.02743  -0.0874   0.7949   0.8966
   0.750   0.2391   0.03322   0.02772  -0.0823   0.7861   0.9079
   1.000   0.2652   0.03295   0.02738  -0.0818   0.7832   0.9165
   1.250   0.2959   0.03267   0.02702  -0.0821   0.7814   0.9259
   1.500   0.3240   0.03215   0.02645  -0.0814   0.7800   0.9368
   1.750   0.3190   0.03248   0.02679  -0.0764   0.7691   0.9476
   2.000   0.3523   0.03206   0.02632  -0.0772   0.7672   0.9545
   2.250   0.3893   0.03172   0.02592  -0.0789   0.7659   0.9600
   2.500   0.3956   0.03236   0.02657  -0.0764   0.7551   0.9668
   2.750   0.4339   0.03207   0.02625  -0.0784   0.7530   0.9717
   3.000   0.4755   0.03175   0.02590  -0.0811   0.7516   0.9761
   3.250   0.5199   0.03135   0.02549  -0.0842   0.7505   0.9803
   3.500   0.5333   0.03210   0.02626  -0.0833   0.7389   0.9913
   3.750   0.5741   0.03171   0.02587  -0.0858   0.7370   1.0000
   4.000   0.6167   0.03127   0.02542  -0.0885   0.7358   1.0000
   4.250   0.6596   0.03081   0.02497  -0.0913   0.7349   1.0000
   4.500   0.7039   0.03018   0.02436  -0.0941   0.7341   1.0000
   4.750   0.7138   0.03105   0.02527  -0.0925   0.7213   1.0000
   5.000   0.7591   0.03020   0.02444  -0.0953   0.7204   1.0000
   5.250   0.8041   0.02931   0.02359  -0.0981   0.7195   1.0000
   5.500   0.8497   0.02831   0.02265  -0.1009   0.7189   1.0000
   5.750   0.8965   0.02717   0.02157  -0.1039   0.7185   1.0000
   7.250   1.0818   0.02493   0.01982  -0.1083   0.6790   1.0000
   7.500   1.1302   0.02348   0.01846  -0.1113   0.6756   1.0000
   7.750   1.1495   0.02354   0.01861  -0.1103   0.6630   1.0000
   8.000   1.1756   0.02328   0.01845  -0.1102   0.6506   1.0000
   8.250   1.2078   0.02274   0.01798  -0.1109   0.6371   1.0000
   8.500   1.2443   0.02202   0.01731  -0.1122   0.6206   1.0000
   8.750   1.2851   0.02117   0.01644  -0.1140   0.5993   1.0000
   9.000   1.3093   0.02123   0.01639  -0.1135   0.5661   1.0000
   9.250   1.3259   0.02177   0.01674  -0.1120   0.5267   1.0000
   9.500   1.3331   0.02286   0.01760  -0.1093   0.4834   1.0000
   9.750   1.3372   0.02422   0.01877  -0.1064   0.4451   1.0000
  10.000   1.3388   0.02581   0.02020  -0.1033   0.4065   1.0000
  10.250   1.3396   0.02757   0.02178  -0.1003   0.3704   1.0000
  10.500   1.3402   0.02944   0.02349  -0.0976   0.3349   1.0000
  10.750   1.3397   0.03151   0.02538  -0.0949   0.2964   1.0000
  11.000   1.3401   0.03363   0.02732  -0.0925   0.2628   1.0000
  11.250   1.3415   0.03579   0.02932  -0.0904   0.2328   1.0000
  11.500   1.3455   0.03785   0.03128  -0.0887   0.2026   1.0000
  11.750   1.3491   0.04002   0.03332  -0.0870   0.1714   1.0000
  12.000   1.3512   0.04237   0.03552  -0.0854   0.1399   1.0000
  12.250   1.3504   0.04506   0.03799  -0.0836   0.1049   1.0000
  12.500   1.3418   0.04858   0.04117  -0.0813   0.0619   1.0000
  12.750   1.3346   0.05215   0.04458  -0.0791   0.0423   1.0000
  13.000   1.3324   0.05534   0.04777  -0.0776   0.0358   1.0000
  13.250   1.3364   0.05794   0.05047  -0.0766   0.0322   1.0000
  13.500   1.3346   0.06121   0.05379  -0.0755   0.0300   1.0000
  13.750   1.3375   0.06405   0.05676  -0.0747   0.0283   1.0000
  14.000   1.3412   0.06683   0.05966  -0.0740   0.0269   1.0000
  14.250   1.3446   0.06968   0.06261  -0.0734   0.0259   1.0000
  14.500   1.3479   0.07258   0.06558  -0.0729   0.0249   1.0000
  14.750   1.3512   0.07547   0.06852  -0.0724   0.0242   1.0000
  15.000   1.3570   0.07799   0.07105  -0.0715   0.0234   1.0000
  15.250   1.3674   0.08010   0.07329  -0.0710   0.0230   1.0000
  15.500   1.3789   0.08214   0.07546  -0.0706   0.0225   1.0000
  15.750   1.3909   0.08415   0.07763  -0.0702   0.0221   1.0000
  16.000   1.4014   0.08638   0.08001  -0.0699   0.0215   1.0000
  16.250   1.4099   0.08885   0.08263  -0.0697   0.0209   1.0000
  16.500   1.4171   0.09146   0.08539  -0.0697   0.0203   1.0000
  16.750   1.4231   0.09424   0.08831  -0.0698   0.0199   1.0000
  17.000   1.4290   0.09713   0.09136  -0.0699   0.0196   1.0000
  17.250   1.4332   0.10029   0.09472  -0.0700   0.0194   1.0000
  17.500   1.4349   0.10384   0.09846  -0.0704   0.0193   1.0000
  17.750   1.4335   0.10781   0.10268  -0.0711   0.0192   1.0000
  18.000   1.4283   0.11235   0.10747  -0.0721   0.0192   1.0000
  18.250   1.4196   0.11747   0.11286  -0.0735   0.0193   1.0000
  18.500   1.4082   0.12315   0.11881  -0.0755   0.0194   1.0000
  18.750   1.3922   0.12977   0.12573  -0.0783   0.0196   1.0000
  19.000   1.3668   0.13839   0.13471  -0.0828   0.0198   1.0000
  19.250   1.3392   0.14801   0.14468  -0.0884   0.0202   1.0000
<< Back to EPPLER 694 AIRFOIL (e694-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 694 AIRFOIL (e694-il)