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OAF095 AIRFOIL (oaf095-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: OAF095 AIRFOIL (oaf095-il)
Reynolds number: 100,000
Max Cl/Cd: 65.67 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-oaf095-il-100000.txt
Download as CSV file: xf-oaf095-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: OAF095 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3803   0.09436   0.08932  -0.0278   1.0000   0.1418
  -8.000  -0.3783   0.09168   0.08669  -0.0291   1.0000   0.1469
  -7.750  -0.4161   0.09070   0.08596  -0.0381   1.0000   0.1508
  -7.500  -0.4032   0.08641   0.08168  -0.0330   1.0000   0.1522
  -7.250  -0.4471   0.06126   0.05651  -0.0612   1.0000   0.0949
  -7.000  -0.4411   0.05993   0.05522  -0.0584   1.0000   0.0958
  -6.750  -0.4185   0.04221   0.03674  -0.0810   1.0000   0.0916
  -6.500  -0.3737   0.03209   0.02514  -0.0947   1.0000   0.0943
  -6.250  -0.3522   0.03029   0.02345  -0.0949   1.0000   0.0987
  -6.000  -0.3231   0.02812   0.02083  -0.0970   1.0000   0.1055
  -5.750  -0.2933   0.02587   0.01824  -0.0990   1.0000   0.1116
  -5.500  -0.2653   0.02476   0.01680  -0.1000   1.0000   0.1209
  -5.250  -0.2391   0.02373   0.01577  -0.1005   1.0000   0.1292
  -5.000  -0.2116   0.02289   0.01478  -0.1013   1.0000   0.1410
  -4.750  -0.1830   0.02214   0.01389  -0.1021   1.0000   0.1544
  -4.500  -0.1558   0.02170   0.01340  -0.1027   1.0000   0.1698
  -4.250  -0.1290   0.02135   0.01303  -0.1032   1.0000   0.1861
  -4.000  -0.1029   0.02115   0.01282  -0.1036   1.0000   0.2033
  -3.750  -0.0772   0.02104   0.01270  -0.1039   1.0000   0.2213
  -3.500  -0.0514   0.02098   0.01259  -0.1043   1.0000   0.2399
  -3.250  -0.0161   0.02089   0.01251  -0.1064   0.9970   0.2609
  -3.000   0.0274   0.02084   0.01250  -0.1099   0.9916   0.2848
  -2.750   0.0710   0.02075   0.01239  -0.1133   0.9857   0.3084
  -2.500   0.1131   0.02072   0.01240  -0.1164   0.9795   0.3312
  -2.250   0.1561   0.02065   0.01232  -0.1197   0.9731   0.3550
  -2.000   0.1963   0.02063   0.01236  -0.1224   0.9662   0.3775
  -1.750   0.2375   0.02058   0.01232  -0.1252   0.9593   0.4024
  -1.500   0.2760   0.02057   0.01239  -0.1275   0.9522   0.4266
  -1.250   0.3154   0.02050   0.01240  -0.1299   0.9450   0.4522
  -1.000   0.3561   0.02045   0.01239  -0.1325   0.9386   0.4798
  -0.750   0.3921   0.02040   0.01246  -0.1342   0.9307   0.5051
  -0.500   0.4388   0.02018   0.01235  -0.1378   0.9265   0.5352
  -0.250   0.4682   0.02023   0.01251  -0.1383   0.9164   0.5610
   0.000   0.5108   0.01998   0.01240  -0.1408   0.9113   0.5935
   0.250   0.5381   0.02006   0.01262  -0.1408   0.9010   0.6242
   0.500   0.5774   0.01976   0.01249  -0.1424   0.8955   0.6662
   0.750   0.6019   0.01974   0.01271  -0.1415   0.8845   0.7120
   1.000   0.6283   0.01930   0.01265  -0.1402   0.8768   0.7957
   1.250   0.6519   0.01902   0.01252  -0.1388   0.8662   1.0000
   1.500   0.6830   0.01927   0.01268  -0.1395   0.8556   1.0000
   1.750   0.7187   0.01918   0.01253  -0.1402   0.8487   1.0000
   2.000   0.7454   0.01948   0.01282  -0.1399   0.8369   1.0000
   2.250   0.7742   0.01959   0.01293  -0.1395   0.8264   1.0000
   2.500   0.8059   0.01934   0.01267  -0.1389   0.8176   1.0000
   2.750   0.8324   0.01931   0.01268  -0.1378   0.8041   1.0000
   3.000   0.8593   0.01916   0.01256  -0.1365   0.7905   1.0000
   3.250   0.8862   0.01894   0.01236  -0.1351   0.7766   1.0000
   3.500   0.9130   0.01865   0.01212  -0.1334   0.7618   1.0000
   3.750   0.9395   0.01830   0.01179  -0.1317   0.7456   1.0000
   4.000   0.9659   0.01790   0.01140  -0.1298   0.7278   1.0000
   4.250   0.9923   0.01749   0.01098  -0.1278   0.7089   1.0000
   4.500   1.0170   0.01732   0.01087  -0.1261   0.6852   1.0000
   4.750   1.0427   0.01704   0.01058  -0.1243   0.6607   1.0000
   5.000   1.0673   0.01689   0.01043  -0.1225   0.6307   1.0000
   5.250   1.0907   0.01687   0.01045  -0.1207   0.5920   1.0000
   5.500   1.1131   0.01695   0.01044  -0.1188   0.5409   1.0000
   5.750   1.1317   0.01745   0.01059  -0.1163   0.4513   1.0000
   6.000   1.1386   0.01951   0.01147  -0.1128   0.2838   1.0000
   6.250   1.1455   0.02216   0.01321  -0.1100   0.1898   1.0000
   6.500   1.1593   0.02408   0.01478  -0.1079   0.1514   1.0000
   6.750   1.1751   0.02583   0.01626  -0.1061   0.1309   1.0000
   7.000   1.1936   0.02752   0.01778  -0.1046   0.1162   1.0000
   7.250   1.2149   0.02899   0.01926  -0.1033   0.1052   1.0000
   7.500   1.2378   0.03076   0.02105  -0.1023   0.0967   1.0000
   7.750   1.2618   0.03272   0.02281  -0.1018   0.0892   1.0000
   8.000   1.2868   0.03460   0.02494  -0.1008   0.0843   1.0000
   8.250   1.3101   0.03636   0.02677  -0.1000   0.0791   1.0000
   8.500   1.3355   0.03924   0.02967  -0.0997   0.0750   1.0000
   8.750   1.3564   0.04161   0.03247  -0.0983   0.0728   1.0000
   9.000   1.3753   0.04419   0.03542  -0.0968   0.0703   1.0000
   9.250   1.3939   0.04653   0.03794  -0.0956   0.0674   1.0000
   9.500   1.4117   0.04949   0.04100  -0.0947   0.0651   1.0000
   9.750   1.4253   0.05418   0.04601  -0.0934   0.0641   1.0000
  10.000   1.4323   0.05891   0.05115  -0.0913   0.0638   1.0000
  10.250   1.4339   0.06279   0.05550  -0.0885   0.0636   1.0000
  10.500   1.4302   0.06671   0.05986  -0.0854   0.0634   1.0000
  10.750   1.4214   0.07071   0.06428  -0.0822   0.0632   1.0000
  11.000   1.4083   0.07490   0.06882  -0.0791   0.0631   1.0000
  11.250   1.3922   0.07921   0.07340  -0.0760   0.0633   1.0000
  11.500   1.3764   0.08397   0.07836  -0.0735   0.0636   1.0000
  11.750   1.3623   0.08830   0.08291  -0.0714   0.0641   1.0000
  12.000   1.2363   0.10073   0.09621  -0.0746   0.0716   1.0000
  12.250   1.2060   0.10945   0.10506  -0.0787   0.0733   1.0000
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