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

EPPLER STF 863-615 AIRFOIL (stf86361-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: EPPLER STF 863-615 AIRFOIL (stf86361-il)
Reynolds number: 200,000
Max Cl/Cd: 54.26 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-stf86361-il-200000.txt
Download as CSV file: xf-stf86361-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER STF 863-615 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  -0.5760   0.09700   0.09409  -0.0394   0.9988   0.0295
  -8.500  -0.5929   0.09212   0.08919  -0.0424   0.9962   0.0293
  -8.250  -0.6129   0.08775   0.08477  -0.0445   0.9915   0.0290
  -8.000  -0.6238   0.08252   0.07940  -0.0486   0.9885   0.0292
  -7.750  -0.6333   0.07459   0.07084  -0.0549   0.9827   0.0302
  -7.500  -0.6224   0.07170   0.06809  -0.0551   0.9813   0.0308
  -7.250  -0.6156   0.06839   0.06468  -0.0554   0.9774   0.0319
  -7.000  -0.5989   0.06322   0.05909  -0.0590   0.9747   0.0345
  -6.750  -0.5768   0.05953   0.05502  -0.0621   0.9734   0.0382
  -6.500  -0.5670   0.05691   0.05238  -0.0613   0.9697   0.0401
  -6.250  -0.5399   0.05334   0.04831  -0.0645   0.9670   0.0470
  -6.000  -0.5128   0.05086   0.04569  -0.0670   0.9655   0.0531
  -5.750  -0.4813   0.04858   0.04310  -0.0704   0.9645   0.0637
  -5.500  -0.4646   0.04641   0.04066  -0.0704   0.9610   0.0757
  -5.250  -0.4366   0.04446   0.03865  -0.0727   0.9580   0.0923
  -4.250  -0.2922   0.03519   0.02735  -0.0744   0.9508   0.0452
  -4.000  -0.2602   0.03399   0.02597  -0.0745   0.9478   0.0400
  -3.750  -0.2251   0.03346   0.02523  -0.0755   0.9459   0.0379
  -3.500  -0.1885   0.03312   0.02500  -0.0778   0.9446   0.0408
  -3.250  -0.1463   0.03322   0.02510  -0.0811   0.9436   0.0416
  -3.000  -0.1290   0.03223   0.02412  -0.0800   0.9359   0.0421
  -2.750  -0.0811   0.03226   0.02414  -0.0850   0.9333   0.0445
  -2.500  -0.0530   0.03203   0.02385  -0.0859   0.9252   0.0496
  -2.250   0.0416   0.03026   0.02430  -0.1003   0.9094   0.6221
  -2.000   0.0613   0.03129   0.02534  -0.0972   0.8993   0.6678
  -1.750   0.0852   0.03227   0.02628  -0.0954   0.8945   0.6929
  -1.500   0.1034   0.03286   0.02691  -0.0924   0.8874   0.7056
  -1.250   0.1197   0.03327   0.02731  -0.0897   0.8814   0.7172
  -1.000   0.1479   0.03387   0.02788  -0.0890   0.8782   0.7327
  -0.750   0.1601   0.03426   0.02828  -0.0855   0.8704   0.7459
  -0.500   0.1802   0.03438   0.02839  -0.0836   0.8650   0.7537
  -0.250   0.2095   0.03469   0.02869  -0.0829   0.8620   0.7693
   0.000   0.2196   0.03487   0.02887  -0.0793   0.8527   0.7820
   0.250   0.2473   0.03477   0.02876  -0.0792   0.8482   0.7878
   0.500   0.2890   0.03474   0.02868  -0.0823   0.8457   0.7886
   0.750   0.3312   0.03479   0.02870  -0.0854   0.8442   0.7900
   1.000   0.3408   0.03479   0.02870  -0.0832   0.8329   0.7912
   1.250   0.3809   0.03470   0.02859  -0.0861   0.8302   0.7918
   1.500   0.4225   0.03463   0.02852  -0.0892   0.8285   0.7925
   1.750   0.4639   0.03464   0.02854  -0.0922   0.8274   0.7938
   2.000   0.4712   0.03480   0.02873  -0.0898   0.8153   0.7947
   2.250   0.5101   0.03473   0.02871  -0.0924   0.8134   0.7952
   2.500   0.5509   0.03461   0.02862  -0.0954   0.8120   0.7957
   2.750   0.5623   0.03491   0.02896  -0.0936   0.8008   0.7964
   3.000   0.6002   0.03474   0.02885  -0.0960   0.7984   0.7975
   3.250   0.6421   0.03438   0.02859  -0.0989   0.7967   0.7984
   3.500   0.6578   0.03457   0.02885  -0.0977   0.7860   0.7989
   3.750   0.6972   0.03414   0.02850  -0.1001   0.7832   0.7994
   4.000   0.7395   0.03353   0.02800  -0.1028   0.7813   0.7999
   4.250   0.7585   0.03350   0.02809  -0.1019   0.7705   0.8003
   4.500   0.8021   0.03249   0.02722  -0.1045   0.7675   0.8009
   4.750   0.8638   0.02969   0.02459  -0.1087   0.7647   0.8018
   5.000   0.9002   0.02722   0.02228  -0.1085   0.7503   0.8028
   5.500   1.0168   0.01874   0.01354  -0.1123   0.6113   0.8043
   6.000   0.9844   0.02272   0.01553  -0.0978   0.3027   0.8053
   6.250   0.9632   0.02567   0.01737  -0.0911   0.1248   0.8058
   6.500   0.9602   0.02782   0.01894  -0.0870   0.0507   0.8063
   6.750   0.9722   0.02896   0.02009  -0.0851   0.0428   0.8070
   7.000   0.9843   0.03009   0.02128  -0.0832   0.0388   0.8078
   7.250   0.9958   0.03124   0.02250  -0.0814   0.0354   0.8086
   7.500   1.0062   0.03252   0.02378  -0.0794   0.0331   0.8093
   7.750   1.0184   0.03373   0.02506  -0.0776   0.0315   0.8103
   8.000   1.0309   0.03509   0.02636  -0.0759   0.0302   0.8115
   8.250   1.0500   0.03622   0.02759  -0.0747   0.0292   0.8128
   8.500   1.0756   0.03730   0.02871  -0.0744   0.0279   0.8140
   8.750   1.1040   0.03844   0.02990  -0.0747   0.0256   0.8151
   9.000   1.1454   0.03983   0.03143  -0.0763   0.0243   0.8162
   9.250   1.1004   0.03486   0.02705  -0.0639   0.0244   0.8173
   9.500   1.2139   0.04185   0.03447  -0.0751   0.0264   0.8186
  13.500   0.9610   0.13812   0.13553  -0.0592   0.0554   0.8312
  13.750   0.9442   0.15014   0.14752  -0.0677   0.0553   0.8319
  14.000   0.9368   0.15817   0.15552  -0.0731   0.0537   0.8330
<< Back to EPPLER STF 863-615 AIRFOIL (stf86361-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER STF 863-615 AIRFOIL (stf86361-il)