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

EPPLER 793 AIRFOIL (e793-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 793 AIRFOIL (e793-il)
Reynolds number: 50,000
Max Cl/Cd: 9.29 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e793-il-50000.txt
Download as CSV file: xf-e793-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 793 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.2452   0.11632   0.10950  -0.0306   1.0000   0.2683
  -9.250  -0.2626   0.11574   0.10906  -0.0302   1.0000   0.2749
  -9.000  -0.2708   0.11347   0.10692  -0.0296   1.0000   0.2769
  -8.750  -0.2562   0.10926   0.10275  -0.0285   1.0000   0.2798
  -8.500  -0.2545   0.10665   0.10022  -0.0273   1.0000   0.2828
  -8.250  -0.2608   0.10468   0.09837  -0.0258   1.0000   0.2869
  -8.000  -0.2781   0.10348   0.09732  -0.0239   1.0000   0.2901
  -7.750  -0.2838   0.10096   0.09490  -0.0221   1.0000   0.2875
  -7.500  -0.5213   0.07333   0.06763  -0.0460   1.0000   0.1339
  -7.250  -0.5330   0.06944   0.06371  -0.0459   1.0000   0.1329
  -7.000  -0.5435   0.06505   0.05925  -0.0467   1.0000   0.1319
  -6.750  -0.5516   0.05977   0.05376  -0.0489   1.0000   0.1307
  -6.500  -0.5524   0.05400   0.04757  -0.0521   1.0000   0.1304
  -6.250  -0.5433   0.04897   0.04199  -0.0550   1.0000   0.1311
  -6.000  -0.5273   0.04459   0.03693  -0.0575   1.0000   0.1327
  -5.750  -0.5103   0.04192   0.03401  -0.0581   1.0000   0.1360
  -5.500  -0.4938   0.04052   0.03257  -0.0578   1.0000   0.1421
  -5.250  -0.4735   0.03834   0.02996  -0.0588   1.0000   0.1504
  -5.000  -0.4552   0.03713   0.02868  -0.0586   1.0000   0.1613
  -4.750  -0.4063   0.03597   0.02759  -0.0634   0.9885   0.1897
  -4.500  -0.3611   0.03577   0.02763  -0.0673   0.9761   0.2389
  -4.250  -0.3218   0.03690   0.02915  -0.0692   0.9626   0.2974
  -4.000  -0.2861   0.03802   0.03045  -0.0700   0.9492   0.3477
  -3.750  -0.2538   0.03895   0.03150  -0.0699   0.9361   0.3876
  -3.500  -0.2212   0.03961   0.03216  -0.0701   0.9231   0.4233
  -3.250  -0.1866   0.04011   0.03262  -0.0706   0.9110   0.4560
  -3.000  -0.1527   0.04027   0.03267  -0.0715   0.8989   0.4863
  -2.750  -0.1285   0.04048   0.03285  -0.0706   0.8858   0.5097
  -2.500  -0.0970   0.04042   0.03261  -0.0719   0.8739   0.5364
  -2.250  -0.0600   0.04051   0.03266  -0.0725   0.8635   0.5598
  -2.000  -0.0382   0.04055   0.03261  -0.0719   0.8506   0.5796
  -1.750  -0.0142   0.04063   0.03260  -0.0718   0.8390   0.5999
  -1.500   0.0270   0.04044   0.03226  -0.0740   0.8292   0.6243
  -1.250   0.0461   0.04060   0.03236  -0.0731   0.8170   0.6421
  -1.000   0.0667   0.04082   0.03253  -0.0724   0.8059   0.6606
  -0.750   0.1046   0.04066   0.03229  -0.0734   0.7969   0.6829
  -0.500   0.1157   0.04114   0.03272  -0.0718   0.7848   0.7004
  -0.250   0.1433   0.04130   0.03281  -0.0720   0.7755   0.7226
   0.000   0.1675   0.04150   0.03296  -0.0717   0.7652   0.7450
   0.250   0.1798   0.04205   0.03350  -0.0700   0.7548   0.7655
   0.500   0.2125   0.04197   0.03336  -0.0702   0.7465   0.7946
   0.750   0.2129   0.04283   0.03427  -0.0673   0.7360   0.8172
   1.000   0.2451   0.04248   0.03391  -0.0666   0.7287   0.8543
   1.250   0.2396   0.04356   0.03509  -0.0635   0.7182   0.8897
   1.500   0.3042   0.04356   0.03514  -0.0703   0.7085   1.0000
   1.750   0.3405   0.04476   0.03611  -0.0763   0.6977   1.0000
   2.000   0.3724   0.04628   0.03741  -0.0809   0.6880   1.0000
   2.250   0.4166   0.04710   0.03799  -0.0849   0.6792   1.0000
   2.500   0.4246   0.04930   0.04004  -0.0854   0.6696   1.0000
   2.750   0.4649   0.05006   0.04063  -0.0878   0.6618   1.0000
   3.000   0.4631   0.05268   0.04315  -0.0871   0.6530   1.0000
   3.250   0.4964   0.05370   0.04404  -0.0884   0.6452   1.0000
   3.500   0.4966   0.05635   0.04662  -0.0878   0.6378   1.0000
   3.750   0.5115   0.05827   0.04847  -0.0879   0.6304   1.0000
   4.000   0.5333   0.05998   0.05010  -0.0884   0.6234   1.0000
   4.250   0.5262   0.06310   0.05319  -0.0876   0.6178   1.0000
   4.500   0.5605   0.06425   0.05427  -0.0886   0.6106   1.0000
   4.750   0.5561   0.06738   0.05737  -0.0880   0.6055   1.0000
   5.000   0.5522   0.07060   0.06059  -0.0878   0.6027   1.0000
   5.250   0.5541   0.07363   0.06361  -0.0879   0.6007   1.0000
   5.500   0.5569   0.07670   0.06668  -0.0882   0.5999   1.0000
   5.750   0.5582   0.08030   0.07028  -0.0888   0.6045   1.0000
   6.000   0.5785   0.08369   0.07367  -0.0906   0.6094   1.0000
   6.250   0.4835   0.09143   0.08160  -0.0911   0.6991   1.0000
   6.500   0.5121   0.09501   0.08515  -0.0930   0.6931   1.0000
   6.750   0.5169   0.09643   0.08658  -0.0922   0.6804   1.0000
   7.000   0.5260   0.09882   0.08896  -0.0922   0.6715   1.0000
   7.250   0.5558   0.10200   0.09213  -0.0938   0.6612   1.0000
   7.500   0.5531   0.10363   0.09379  -0.0928   0.6508   1.0000
   7.750   0.5938   0.10806   0.09822  -0.0953   0.6430   1.0000
   8.000   0.5855   0.10889   0.09907  -0.0938   0.6302   1.0000
   8.250   0.5990   0.11192   0.10211  -0.0943   0.6229   1.0000
   8.500   0.6179   0.11456   0.10478  -0.0949   0.6113   1.0000
   8.750   0.6188   0.11668   0.10693  -0.0945   0.6008   1.0000
   9.000   0.6560   0.12106   0.11134  -0.0964   0.5926   1.0000
   9.250   0.6456   0.12215   0.11247  -0.0954   0.5808   1.0000
   9.500   0.6684   0.12605   0.11641  -0.0965   0.5738   1.0000
<< Back to EPPLER 793 AIRFOIL (e793-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 793 AIRFOIL (e793-il)