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EPPLER STF 863-615 AIRFOIL (stf86361-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER STF 863-615 AIRFOIL (stf86361-il)
Reynolds number: 100,000
Max Cl/Cd: 32.11 at α=7.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-stf86361-il-100000.txt
Download as CSV file: xf-stf86361-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5305   0.10749   0.10376  -0.0310   1.0019   0.1166
  -8.250  -0.5138   0.10562   0.10189  -0.0270   1.0019   0.1254
  -8.000  -0.5413   0.10076   0.09708  -0.0282   1.0019   0.1266
  -7.750  -0.5697   0.09623   0.09261  -0.0282   1.0019   0.1266
  -7.500  -0.6569   0.09031   0.08640  -0.0325   1.0019   0.1124
  -7.250  -0.6823   0.08459   0.08030  -0.0387   1.0019   0.1168
  -7.000  -0.6697   0.08159   0.07753  -0.0350   1.0019   0.1235
  -6.750  -0.6709   0.07705   0.07286  -0.0368   1.0019   0.1325
  -6.500  -0.6666   0.07310   0.06880  -0.0380   1.0019   0.1446
  -6.250  -0.6586   0.06973   0.06537  -0.0382   1.0019   0.1596
  -6.000  -0.6486   0.06665   0.06224  -0.0383   1.0019   0.1784
  -5.750  -0.6394   0.06343   0.05900  -0.0386   1.0019   0.2073
  -5.500  -0.6308   0.06121   0.05688  -0.0367   1.0019   0.2439
  -4.500  -0.4555   0.04028   0.03190  -0.0529   1.0019   0.0804
  -4.250  -0.4246   0.03815   0.02915  -0.0522   1.0019   0.0674
  -4.000  -0.3965   0.03668   0.02724  -0.0516   1.0019   0.0618
  -3.750  -0.3712   0.03541   0.02589  -0.0511   1.0019   0.0611
  -3.500  -0.3459   0.03448   0.02491  -0.0508   1.0019   0.0634
  -3.250  -0.3205   0.03379   0.02416  -0.0502   1.0019   0.0633
  -3.000  -0.2947   0.03326   0.02365  -0.0498   1.0019   0.0635
  -2.750  -0.2677   0.03290   0.02335  -0.0501   1.0019   0.0649
  -2.500  -0.2393   0.03275   0.02320  -0.0509   1.0019   0.0689
  -2.250  -0.2097   0.03273   0.02316  -0.0522   1.0019   0.0754
  -2.000  -0.1791   0.03282   0.02319  -0.0537   1.0019   0.0834
  -1.750  -0.1331   0.03289   0.02573  -0.0582   1.0019   0.6570
  -1.500  -0.1372   0.03430   0.02725  -0.0499   1.0019   0.7056
  -1.250  -0.1397   0.03525   0.02826  -0.0420   1.0019   0.7386
  -1.000  -0.1382   0.03598   0.02898  -0.0355   1.0019   0.7687
  -0.750  -0.1280   0.03660   0.02954  -0.0315   1.0005   0.7881
  -0.500  -0.1110   0.03754   0.03040  -0.0286   0.9973   0.8090
  -0.250  -0.0985   0.03807   0.03085  -0.0253   0.9930   0.8297
   0.000  -0.0741   0.03911   0.03179  -0.0247   0.9895   0.8453
   0.250  -0.0593   0.03936   0.03198  -0.0226   0.9834   0.8604
   0.500  -0.0457   0.03967   0.03225  -0.0199   0.9767   0.8778
   0.750  -0.0330   0.03992   0.03246  -0.0171   0.9694   0.8957
   1.000   0.0030   0.04127   0.03367  -0.0200   0.9642   0.9040
   1.250   0.0287   0.04161   0.03395  -0.0215   0.9563   0.9043
   1.500   0.0542   0.04220   0.03450  -0.0229   0.9495   0.9053
   1.750   0.0973   0.04382   0.03606  -0.0275   0.9435   0.9064
   2.000   0.1198   0.04402   0.03624  -0.0283   0.9341   0.9068
   2.250   0.1459   0.04472   0.03693  -0.0299   0.9261   0.9072
   2.500   0.1902   0.04629   0.03848  -0.0347   0.9187   0.9079
   2.750   0.2256   0.04696   0.03915  -0.0374   0.9050   0.9090
   3.000   0.2785   0.04589   0.03806  -0.0411   0.8652   0.9100
   3.250   0.3148   0.04619   0.03839  -0.0436   0.8513   0.9107
   3.500   0.3632   0.04694   0.03921  -0.0482   0.8445   0.9113
   3.750   0.3843   0.04714   0.03946  -0.0485   0.8319   0.9120
   4.000   0.4163   0.04753   0.03993  -0.0504   0.8213   0.9128
   4.250   0.4604   0.04779   0.04027  -0.0540   0.8130   0.9140
   4.500   0.4887   0.04780   0.04038  -0.0550   0.7998   0.9155
   4.750   0.5220   0.04782   0.04056  -0.0568   0.7881   0.9170
   5.000   0.5691   0.04778   0.04067  -0.0605   0.7814   0.9182
   5.250   0.5942   0.04796   0.04099  -0.0612   0.7690   0.9191
   5.750   0.6721   0.04730   0.04074  -0.0656   0.7509   0.9217
   6.000   0.7007   0.04695   0.04059  -0.0661   0.7378   0.9233
   6.250   0.7359   0.04624   0.04011  -0.0672   0.7263   0.9250
   6.500   0.7917   0.04399   0.03821  -0.0701   0.7182   0.9271
   6.750   0.8805   0.03340   0.02801  -0.0696   0.6826   0.9301
   7.000   0.9108   0.02996   0.02474  -0.0664   0.6279   0.9333
   7.250   0.9212   0.02869   0.02072  -0.0588   0.2102   0.9353
   7.500   0.9107   0.03170   0.02255  -0.0541   0.0885   0.9376
   7.750   0.9175   0.03335   0.02413  -0.0516   0.0721   0.9405
   8.000   0.9262   0.03489   0.02571  -0.0495   0.0651   0.9437
   8.250   0.9367   0.03640   0.02724  -0.0477   0.0608   0.9474
   8.500   0.9550   0.03781   0.02853  -0.0467   0.0569   0.9517
   8.750   0.9863   0.03899   0.02965  -0.0472   0.0514   0.9576
   9.000   1.0398   0.04037   0.03112  -0.0500   0.0478   0.9663
   9.250   1.1426   0.04433   0.03534  -0.0595   0.0461   0.9804
   9.500   1.1720   0.04727   0.03881  -0.0600   0.0438   0.9981
   9.750   1.2035   0.05145   0.04359  -0.0606   0.0440   0.9981
  10.000   1.2450   0.05738   0.04999  -0.0628   0.0469   0.9981
  10.250   1.2368   0.06175   0.05522  -0.0571   0.0527   0.9981
  10.500   1.0857   0.07463   0.07082  -0.0294   0.1562   0.9981
  10.750   1.0499   0.07852   0.07493  -0.0251   0.1565   0.9981
  11.000   1.0144   0.08302   0.07959  -0.0217   0.1565   0.9981
  11.250   0.9803   0.08798   0.08468  -0.0194   0.1563   0.9981
  11.500   0.9453   0.09342   0.09024  -0.0180   0.1560   0.9981
  11.750   0.9071   0.09934   0.09625  -0.0176   0.1557   0.9981
  12.000   0.8636   0.10595   0.10294  -0.0185   0.1555   0.9981
  12.250   0.8164   0.11362   0.11064  -0.0212   0.1553   0.9981
  12.500   0.7765   0.12199   0.11897  -0.0252   0.1536   0.9981
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