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

OAF128 AIRFOIL (oaf128-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: OAF128 AIRFOIL (oaf128-il)
Reynolds number: 100,000
Max Cl/Cd: 43.54 at α=6.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-oaf128-il-100000-n5.txt
Download as CSV file: xf-oaf128-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: OAF128 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.000  -0.9210   0.09403   0.08826   0.0125   1.0000   0.0410
 -12.750  -0.9909   0.07598   0.06976  -0.0013   1.0000   0.0401
 -12.500  -1.0155   0.06763   0.06116  -0.0077   1.0000   0.0401
 -12.250  -1.0303   0.06134   0.05464  -0.0126   1.0000   0.0403
 -12.000  -1.0405   0.05616   0.04925  -0.0164   1.0000   0.0407
 -11.750  -1.0482   0.05187   0.04473  -0.0191   1.0000   0.0411
 -11.500  -1.0525   0.04842   0.04105  -0.0204   1.0000   0.0417
 -11.250  -1.0488   0.04541   0.03776  -0.0210   1.0000   0.0424
 -11.000  -1.0404   0.04268   0.03473  -0.0214   1.0000   0.0432
 -10.750  -1.0264   0.04062   0.03259  -0.0214   1.0000   0.0440
 -10.500  -1.0103   0.03882   0.03076  -0.0214   1.0000   0.0450
 -10.250  -0.9930   0.03707   0.02888  -0.0216   1.0000   0.0463
 -10.000  -0.9745   0.03530   0.02693  -0.0217   1.0000   0.0481
  -9.750  -0.9546   0.03366   0.02517  -0.0218   1.0000   0.0499
  -9.500  -0.9336   0.03224   0.02372  -0.0220   1.0000   0.0517
  -9.250  -0.9114   0.03081   0.02218  -0.0221   1.0000   0.0539
  -9.000  -0.8885   0.02942   0.02069  -0.0223   1.0000   0.0563
  -8.750  -0.8647   0.02819   0.01945  -0.0226   1.0000   0.0594
  -8.500  -0.8400   0.02698   0.01815  -0.0229   1.0000   0.0634
  -8.250  -0.8146   0.02586   0.01700  -0.0233   1.0000   0.0680
  -8.000  -0.7886   0.02476   0.01589  -0.0238   1.0000   0.0737
  -7.750  -0.7619   0.02375   0.01485  -0.0243   1.0000   0.0816
  -7.500  -0.7346   0.02281   0.01391  -0.0248   1.0000   0.0915
  -7.250  -0.7068   0.02194   0.01305  -0.0253   1.0000   0.1037
  -7.000  -0.6786   0.02116   0.01225  -0.0259   1.0000   0.1171
  -6.750  -0.6501   0.02046   0.01156  -0.0264   1.0000   0.1314
  -6.500  -0.6212   0.01985   0.01098  -0.0270   1.0000   0.1460
  -6.250  -0.5920   0.01932   0.01046  -0.0275   1.0000   0.1613
  -6.000  -0.5626   0.01885   0.00998  -0.0281   1.0000   0.1777
  -5.750  -0.5330   0.01840   0.00951  -0.0286   1.0000   0.1952
  -5.500  -0.5034   0.01792   0.00907  -0.0292   1.0000   0.2127
  -5.250  -0.4737   0.01747   0.00870  -0.0298   1.0000   0.2290
  -5.000  -0.4438   0.01707   0.00836  -0.0303   1.0000   0.2449
  -4.750  -0.4137   0.01671   0.00804  -0.0308   1.0000   0.2611
  -4.500  -0.3835   0.01637   0.00777  -0.0314   1.0000   0.2772
  -4.250  -0.3532   0.01607   0.00752  -0.0320   1.0000   0.2936
  -4.000  -0.3227   0.01579   0.00731  -0.0327   1.0000   0.3100
  -3.750  -0.2897   0.01562   0.00717  -0.0338   0.9734   0.3275
  -3.500  -0.2534   0.01555   0.00710  -0.0353   0.9288   0.3455
  -3.250  -0.2238   0.01553   0.00705  -0.0352   0.8930   0.3620
  -3.000  -0.1972   0.01552   0.00699  -0.0344   0.8615   0.3777
  -2.750  -0.1713   0.01551   0.00693  -0.0334   0.8330   0.3935
  -2.500  -0.1449   0.01549   0.00685  -0.0326   0.8059   0.4099
  -2.250  -0.1183   0.01547   0.00676  -0.0318   0.7811   0.4271
  -2.000  -0.0910   0.01545   0.00668  -0.0312   0.7570   0.4444
  -1.750  -0.0633   0.01541   0.00660  -0.0307   0.7344   0.4601
  -1.500  -0.0355   0.01536   0.00652  -0.0303   0.7131   0.4744
  -1.250  -0.0074   0.01532   0.00644  -0.0299   0.6930   0.4889
  -1.000   0.0210   0.01530   0.00636  -0.0296   0.6742   0.5036
  -0.750   0.0495   0.01529   0.00630  -0.0293   0.6565   0.5190
  -0.500   0.0779   0.01528   0.00628  -0.0291   0.6398   0.5348
  -0.250   0.1064   0.01527   0.00626  -0.0288   0.6235   0.5516
   0.000   0.1348   0.01527   0.00624  -0.0285   0.6068   0.5700
   0.250   0.1631   0.01528   0.00624  -0.0282   0.5897   0.5907
   0.500   0.1915   0.01528   0.00627  -0.0280   0.5729   0.6146
   0.750   0.2196   0.01527   0.00632  -0.0276   0.5578   0.6399
   1.000   0.2476   0.01528   0.00639  -0.0272   0.5444   0.6677
   1.250   0.2749   0.01530   0.00644  -0.0266   0.5320   0.6970
   1.500   0.3022   0.01530   0.00654  -0.0261   0.5184   0.7298
   1.750   0.3282   0.01530   0.00662  -0.0252   0.5062   0.7665
   2.250   0.3764   0.01522   0.00671  -0.0223   0.4825   0.8557
   2.500   0.4009   0.01510   0.00665  -0.0210   0.4709   0.9203
   2.750   0.4330   0.01513   0.00661  -0.0216   0.4585   1.0000
   3.000   0.4641   0.01535   0.00679  -0.0223   0.4453   1.0000
   3.250   0.4946   0.01560   0.00696  -0.0228   0.4326   1.0000
   3.500   0.5247   0.01585   0.00711  -0.0231   0.4202   1.0000
   3.750   0.5548   0.01609   0.00732  -0.0235   0.4053   1.0000
   4.000   0.5847   0.01633   0.00749  -0.0239   0.3882   1.0000
   4.250   0.6142   0.01657   0.00767  -0.0242   0.3703   1.0000
   4.500   0.6437   0.01684   0.00788  -0.0244   0.3530   1.0000
   4.750   0.6731   0.01713   0.00813  -0.0247   0.3364   1.0000
   5.000   0.7023   0.01744   0.00841  -0.0250   0.3200   1.0000
   5.250   0.7313   0.01778   0.00871  -0.0252   0.3035   1.0000
   5.500   0.7601   0.01815   0.00904  -0.0255   0.2868   1.0000
   5.750   0.7887   0.01854   0.00941  -0.0257   0.2697   1.0000
   6.000   0.8169   0.01898   0.00980  -0.0259   0.2522   1.0000
   6.250   0.8449   0.01948   0.01024  -0.0261   0.2348   1.0000
   6.500   0.8723   0.02004   0.01075  -0.0263   0.2173   1.0000
   6.750   0.8995   0.02066   0.01131  -0.0265   0.2001   1.0000
   7.000   0.9265   0.02130   0.01194  -0.0266   0.1829   1.0000
   7.250   0.9530   0.02200   0.01262  -0.0267   0.1671   1.0000
   7.500   0.9788   0.02277   0.01333  -0.0268   0.1537   1.0000
   7.750   1.0042   0.02358   0.01412  -0.0268   0.1424   1.0000
   8.000   1.0293   0.02442   0.01498  -0.0267   0.1332   1.0000
   8.250   1.0533   0.02536   0.01588  -0.0266   0.1259   1.0000
   8.500   1.0773   0.02627   0.01684  -0.0265   0.1192   1.0000
   8.750   1.0999   0.02732   0.01785  -0.0262   0.1141   1.0000
   9.000   1.1229   0.02829   0.01892  -0.0259   0.1089   1.0000
   9.250   1.1444   0.02937   0.01998  -0.0256   0.1047   1.0000
   9.500   1.1657   0.03050   0.02118  -0.0251   0.1009   1.0000
   9.750   1.1867   0.03162   0.02240  -0.0246   0.0971   1.0000
  10.000   1.2063   0.03279   0.02359  -0.0241   0.0939   1.0000
  10.250   1.2252   0.03409   0.02494  -0.0235   0.0911   1.0000
  10.500   1.2438   0.03539   0.02641  -0.0229   0.0881   1.0000
  10.750   1.2610   0.03672   0.02781  -0.0222   0.0854   1.0000
  11.000   1.2774   0.03807   0.02915  -0.0215   0.0832   1.0000
  11.250   1.2922   0.03964   0.03085  -0.0207   0.0810   1.0000
  11.500   1.3048   0.04131   0.03274  -0.0198   0.0788   1.0000
  11.750   1.3158   0.04298   0.03454  -0.0188   0.0767   1.0000
  12.000   1.3255   0.04462   0.03624  -0.0178   0.0750   1.0000
  12.250   1.3347   0.04623   0.03784  -0.0167   0.0735   1.0000
  12.500   1.3359   0.04852   0.04034  -0.0156   0.0722   1.0000
  12.750   1.3342   0.05128   0.04337  -0.0151   0.0708   1.0000
  13.000   1.3324   0.05426   0.04655  -0.0152   0.0694   1.0000
  13.250   1.3307   0.05737   0.04984  -0.0157   0.0681   1.0000
  13.500   1.3299   0.06047   0.05306  -0.0164   0.0670   1.0000
  14.000   1.3363   0.06597   0.05862  -0.0172   0.0649   1.0000
  14.250   1.3182   0.07180   0.06478  -0.0200   0.0642   1.0000
  14.500   1.2966   0.07848   0.07178  -0.0236   0.0635   1.0000
  14.750   1.2700   0.08636   0.07995  -0.0282   0.0630   1.0000
  15.000   1.2334   0.09644   0.09033  -0.0345   0.0627   1.0000
  15.250   1.1503   0.11697   0.11130  -0.0483   0.0631   1.0000
<< Back to OAF128 AIRFOIL (oaf128-il)

Polar data table (+)

Polar graphs


<< Back to OAF128 AIRFOIL (oaf128-il)