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=50,000 Ncrit=5


Details Polar file
Airfoil: OAF128 AIRFOIL (oaf128-il)
Reynolds number: 50,000
Max Cl/Cd: 31.08 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-oaf128-il-50000-n5.txt
Download as CSV file: xf-oaf128-il-50000-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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.7864   0.08688   0.07931   0.0045   1.0000   0.0756
 -10.500  -0.8701   0.06649   0.05852  -0.0132   1.0000   0.0732
 -10.250  -0.8916   0.05996   0.05171  -0.0177   1.0000   0.0737
 -10.000  -0.8965   0.05521   0.04664  -0.0200   1.0000   0.0753
  -9.750  -0.8955   0.05072   0.04174  -0.0217   1.0000   0.0773
  -9.500  -0.8858   0.04723   0.03792  -0.0226   1.0000   0.0797
  -9.250  -0.8683   0.04535   0.03603  -0.0227   1.0000   0.0830
  -9.000  -0.8529   0.04247   0.03276  -0.0234   1.0000   0.0875
  -8.750  -0.8331   0.04058   0.03085  -0.0236   1.0000   0.0917
  -8.500  -0.8130   0.03836   0.02836  -0.0240   1.0000   0.0980
  -8.250  -0.7913   0.03676   0.02678  -0.0242   1.0000   0.1045
  -8.000  -0.7688   0.03492   0.02479  -0.0245   1.0000   0.1132
  -7.750  -0.7453   0.03327   0.02300  -0.0249   1.0000   0.1240
  -7.500  -0.7209   0.03189   0.02156  -0.0253   1.0000   0.1365
  -7.250  -0.6958   0.03071   0.02036  -0.0257   1.0000   0.1505
  -7.000  -0.6699   0.02971   0.01936  -0.0262   1.0000   0.1659
  -6.750  -0.6435   0.02882   0.01847  -0.0266   1.0000   0.1826
  -6.500  -0.6167   0.02798   0.01762  -0.0270   1.0000   0.2002
  -6.250  -0.5895   0.02722   0.01688  -0.0274   1.0000   0.2178
  -6.000  -0.5620   0.02654   0.01623  -0.0277   1.0000   0.2349
  -5.750  -0.5340   0.02588   0.01556  -0.0280   1.0000   0.2517
  -5.500  -0.5057   0.02524   0.01491  -0.0282   1.0000   0.2685
  -5.250  -0.4772   0.02465   0.01431  -0.0285   1.0000   0.2853
  -5.000  -0.4486   0.02409   0.01374  -0.0288   1.0000   0.3022
  -4.750  -0.4198   0.02359   0.01324  -0.0291   1.0000   0.3195
  -4.500  -0.3909   0.02311   0.01277  -0.0294   1.0000   0.3369
  -4.250  -0.3621   0.02268   0.01236  -0.0297   1.0000   0.3548
  -4.000  -0.3332   0.02227   0.01198  -0.0301   1.0000   0.3728
  -3.750  -0.3047   0.02191   0.01165  -0.0303   1.0000   0.3911
  -3.500  -0.2762   0.02158   0.01138  -0.0306   1.0000   0.4097
  -3.250  -0.2481   0.02128   0.01116  -0.0310   1.0000   0.4287
  -3.000  -0.2208   0.02103   0.01099  -0.0312   1.0000   0.4478
  -2.750  -0.1953   0.02082   0.01089  -0.0314   1.0000   0.4662
  -2.500  -0.1524   0.02069   0.01078  -0.0345   0.9617   0.4879
  -2.250  -0.1106   0.02055   0.01063  -0.0370   0.9287   0.5082
  -2.000  -0.0751   0.02047   0.01053  -0.0380   0.8970   0.5272
  -1.750  -0.0450   0.02044   0.01046  -0.0378   0.8672   0.5455
  -1.500  -0.0180   0.02043   0.01041  -0.0370   0.8395   0.5634
  -1.250   0.0076   0.02042   0.01036  -0.0358   0.8140   0.5820
  -1.000   0.0327   0.02040   0.01034  -0.0345   0.7897   0.6021
  -0.750   0.0574   0.02039   0.01033  -0.0331   0.7675   0.6238
  -0.500   0.0820   0.02037   0.01032  -0.0317   0.7467   0.6474
  -0.250   0.1071   0.02036   0.01030  -0.0305   0.7271   0.6723
   0.000   0.1312   0.02033   0.01032  -0.0290   0.7086   0.6963
   0.250   0.1556   0.02031   0.01033  -0.0275   0.6908   0.7226
   0.500   0.1795   0.02028   0.01034  -0.0260   0.6738   0.7520
   0.750   0.2027   0.02024   0.01034  -0.0242   0.6575   0.7859
   1.000   0.2245   0.02015   0.01030  -0.0221   0.6416   0.8250
   1.250   0.2474   0.02001   0.01017  -0.0202   0.6260   0.8746
   1.500   0.2794   0.01986   0.01002  -0.0203   0.6086   0.9481
   1.750   0.3143   0.01997   0.01000  -0.0219   0.5909   1.0000
   2.000   0.3455   0.02027   0.01016  -0.0229   0.5746   1.0000
   2.250   0.3761   0.02059   0.01036  -0.0235   0.5592   1.0000
   2.500   0.4061   0.02092   0.01058  -0.0239   0.5445   1.0000
   2.750   0.4355   0.02125   0.01078  -0.0241   0.5311   1.0000
   3.000   0.4651   0.02162   0.01110  -0.0245   0.5161   1.0000
   3.250   0.4947   0.02202   0.01147  -0.0248   0.5012   1.0000
   3.500   0.5239   0.02242   0.01183  -0.0251   0.4871   1.0000
   3.750   0.5527   0.02281   0.01215  -0.0251   0.4734   1.0000
   4.000   0.5815   0.02321   0.01252  -0.0251   0.4591   1.0000
   4.250   0.6102   0.02366   0.01298  -0.0253   0.4435   1.0000
   4.500   0.6386   0.02409   0.01342  -0.0254   0.4277   1.0000
   4.750   0.6665   0.02449   0.01380  -0.0253   0.4109   1.0000
   5.000   0.6942   0.02486   0.01414  -0.0252   0.3932   1.0000
   5.250   0.7216   0.02523   0.01450  -0.0251   0.3750   1.0000
   5.500   0.7488   0.02562   0.01489  -0.0250   0.3566   1.0000
   5.750   0.7758   0.02606   0.01532  -0.0249   0.3384   1.0000
   6.000   0.8025   0.02654   0.01580  -0.0247   0.3202   1.0000
   6.250   0.8288   0.02706   0.01631  -0.0246   0.3019   1.0000
   6.500   0.8546   0.02765   0.01689  -0.0245   0.2835   1.0000
   6.750   0.8799   0.02832   0.01752  -0.0243   0.2652   1.0000
   7.000   0.9047   0.02911   0.01829  -0.0241   0.2471   1.0000
   7.250   0.9288   0.03003   0.01920  -0.0240   0.2297   1.0000
   7.500   0.9522   0.03105   0.02017  -0.0237   0.2142   1.0000
   7.750   0.9750   0.03215   0.02124  -0.0235   0.2001   1.0000
   8.000   0.9974   0.03341   0.02259  -0.0232   0.1870   1.0000
   8.250   1.0192   0.03473   0.02400  -0.0230   0.1759   1.0000
   8.500   1.0405   0.03600   0.02522  -0.0226   0.1671   1.0000
   8.750   1.0610   0.03744   0.02679  -0.0222   0.1586   1.0000
   9.000   1.0810   0.03891   0.02831  -0.0218   0.1517   1.0000
   9.250   1.1000   0.04049   0.03002  -0.0213   0.1451   1.0000
   9.500   1.1194   0.04197   0.03146  -0.0208   0.1399   1.0000
   9.750   1.1349   0.04403   0.03382  -0.0203   0.1343   1.0000
  10.000   1.1523   0.04563   0.03545  -0.0198   0.1298   1.0000
  10.250   1.1674   0.04767   0.03760  -0.0192   0.1259   1.0000
  10.500   1.1762   0.05030   0.04056  -0.0187   0.1219   1.0000
  10.750   1.1873   0.05245   0.04284  -0.0180   0.1184   1.0000
  11.000   1.2054   0.05400   0.04429  -0.0174   0.1154   1.0000
  11.250   1.1993   0.05783   0.04854  -0.0169   0.1130   1.0000
  11.500   1.1853   0.06194   0.05299  -0.0164   0.1111   1.0000
  11.750   1.1660   0.06693   0.05827  -0.0174   0.1097   1.0000
  12.000   1.1386   0.07377   0.06539  -0.0205   0.1086   1.0000
  12.250   1.0564   0.09130   0.08334  -0.0325   0.1095   1.0000
<< Back to OAF128 AIRFOIL (oaf128-il)

Polar data table (+)

Polar graphs


<< Back to OAF128 AIRFOIL (oaf128-il)