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

ARA-D 6% AIRFOIL (arad6-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: ARA-D 6% AIRFOIL (arad6-il)
Reynolds number: 1,000,000
Max Cl/Cd: 139.66 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-arad6-il-1000000.txt
Download as CSV file: xf-arad6-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: ARA-D 6% AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3543   0.10592   0.10426  -0.0266   1.0000   0.0096
  -8.750  -0.3577   0.10421   0.10259  -0.0249   1.0000   0.0097
  -8.500  -0.3599   0.10233   0.10073  -0.0236   0.9998   0.0099
  -8.250  -0.3462   0.09874   0.09714  -0.0270   0.9985   0.0102
  -8.000  -0.3322   0.09501   0.09341  -0.0308   0.9968   0.0106
  -7.750  -0.3174   0.09113   0.08954  -0.0350   0.9951   0.0111
  -7.500  -0.2988   0.08620   0.08461  -0.0422   0.9915   0.0120
  -7.250  -0.2759   0.08140   0.07979  -0.0498   0.9888   0.0122
  -7.000  -0.2502   0.07633   0.07471  -0.0577   0.9868   0.0122
  -6.750  -0.2267   0.06970   0.06806  -0.0664   0.9854   0.0126
  -6.500  -0.2008   0.06659   0.06494  -0.0712   0.9841   0.0129
  -6.250  -0.1759   0.06338   0.06171  -0.0760   0.9806   0.0133
  -6.000  -0.1480   0.05974   0.05804  -0.0818   0.9767   0.0138
  -5.750  -0.1166   0.05562   0.05388  -0.0886   0.9732   0.0146
  -5.500  -0.0729   0.04990   0.04807  -0.0988   0.9695   0.0163
  -5.250  -0.0390   0.04496   0.04305  -0.1053   0.9640   0.0164
  -5.000   0.0007   0.03645   0.03434  -0.1156   0.9596   0.0172
  -4.750   0.0644   0.01874   0.01561  -0.1313   0.9578   0.0167
  -4.500   0.0997   0.01382   0.01000  -0.1345   0.9541   0.0151
  -4.250   0.1301   0.01208   0.00791  -0.1352   0.9492   0.0150
  -4.000   0.1589   0.01122   0.00685  -0.1354   0.9444   0.0153
  -3.750   0.1880   0.01039   0.00584  -0.1356   0.9395   0.0155
  -3.500   0.2171   0.00966   0.00496  -0.1357   0.9339   0.0158
  -3.250   0.2455   0.00919   0.00438  -0.1356   0.9287   0.0162
  -3.000   0.2745   0.00870   0.00380  -0.1358   0.9228   0.0167
  -2.750   0.3036   0.00815   0.00315  -0.1359   0.9167   0.0179
  -2.500   0.3322   0.00791   0.00285  -0.1359   0.9107   0.0189
  -2.250   0.3607   0.00770   0.00261  -0.1359   0.9037   0.0201
  -2.000   0.3892   0.00755   0.00241  -0.1359   0.8972   0.0216
  -1.750   0.4181   0.00732   0.00217  -0.1359   0.8897   0.0279
  -1.500   0.4468   0.00713   0.00204  -0.1360   0.8827   0.0513
  -1.250   0.4751   0.00704   0.00197  -0.1360   0.8738   0.0642
  -1.000   0.5030   0.00698   0.00189  -0.1358   0.8599   0.0778
  -0.750   0.5307   0.00693   0.00183  -0.1356   0.8436   0.0954
  -0.500   0.5587   0.00690   0.00180  -0.1356   0.8292   0.1154
  -0.250   0.5867   0.00689   0.00177  -0.1355   0.8138   0.1333
   0.000   0.6148   0.00689   0.00176  -0.1354   0.7979   0.1526
   0.250   0.6429   0.00690   0.00175  -0.1354   0.7822   0.1705
   0.500   0.6711   0.00691   0.00176  -0.1354   0.7671   0.1913
   0.750   0.6993   0.00692   0.00179  -0.1355   0.7507   0.2218
   1.000   0.7275   0.00688   0.00184  -0.1356   0.7330   0.2836
   1.250   0.7496   0.00545   0.00190  -0.1347   0.7161   1.0000
   1.500   0.7773   0.00561   0.00194  -0.1345   0.6953   1.0000
   1.750   0.8050   0.00577   0.00199  -0.1344   0.6727   1.0000
   2.000   0.8324   0.00596   0.00206  -0.1343   0.6466   1.0000
   2.250   0.8596   0.00618   0.00215  -0.1341   0.6184   1.0000
   2.500   0.8865   0.00645   0.00225  -0.1340   0.5837   1.0000
   2.750   0.9132   0.00675   0.00239  -0.1337   0.5453   1.0000
   3.000   0.9394   0.00713   0.00255  -0.1335   0.4982   1.0000
   3.250   0.9646   0.00766   0.00277  -0.1331   0.4340   1.0000
   3.500   0.9892   0.00829   0.00305  -0.1327   0.3607   1.0000
   3.750   1.0137   0.00893   0.00336  -0.1323   0.2928   1.0000
   4.000   1.0393   0.00937   0.00361  -0.1320   0.2541   1.0000
   4.250   1.0653   0.00974   0.00384  -0.1317   0.2261   1.0000
   4.500   1.0913   0.01009   0.00407  -0.1315   0.2037   1.0000
   4.750   1.1177   0.01038   0.00430  -0.1313   0.1879   1.0000
   5.000   1.1438   0.01067   0.00452  -0.1311   0.1729   1.0000
   5.250   1.1699   0.01096   0.00476  -0.1308   0.1590   1.0000
   5.500   1.1958   0.01126   0.00500  -0.1306   0.1466   1.0000
   5.750   1.2218   0.01154   0.00525  -0.1303   0.1364   1.0000
   6.000   1.2476   0.01182   0.00550  -0.1300   0.1260   1.0000
   6.250   1.2731   0.01214   0.00578  -0.1297   0.1152   1.0000
   6.500   1.2983   0.01248   0.00607  -0.1293   0.1049   1.0000
   6.750   1.3231   0.01284   0.00640  -0.1289   0.0944   1.0000
   7.000   1.3478   0.01321   0.00672  -0.1284   0.0841   1.0000
   7.250   1.3723   0.01359   0.00707  -0.1279   0.0746   1.0000
   7.500   1.3963   0.01400   0.00744  -0.1273   0.0657   1.0000
   7.750   1.4197   0.01446   0.00787  -0.1267   0.0567   1.0000
   8.000   1.4436   0.01485   0.00826  -0.1261   0.0511   1.0000
   8.250   1.4668   0.01531   0.00869  -0.1254   0.0446   1.0000
   8.500   1.4895   0.01578   0.00915  -0.1246   0.0393   1.0000
   8.750   1.5123   0.01623   0.00962  -0.1239   0.0349   1.0000
   9.000   1.5346   0.01671   0.01010  -0.1230   0.0310   1.0000
   9.250   1.5564   0.01723   0.01063  -0.1221   0.0268   1.0000
   9.500   1.5774   0.01779   0.01118  -0.1211   0.0224   1.0000
   9.750   1.5970   0.01848   0.01184  -0.1198   0.0164   1.0000
  10.000   1.6143   0.01939   0.01272  -0.1181   0.0105   1.0000
  10.250   1.6318   0.02021   0.01358  -0.1165   0.0085   1.0000
  10.500   1.6490   0.02102   0.01445  -0.1147   0.0075   1.0000
  10.750   1.6664   0.02173   0.01524  -0.1131   0.0069   1.0000
  11.000   1.6820   0.02254   0.01613  -0.1111   0.0064   1.0000
  11.250   1.6954   0.02350   0.01717  -0.1088   0.0059   1.0000
  11.500   1.7039   0.02480   0.01860  -0.1057   0.0055   1.0000
  11.750   1.7126   0.02585   0.01977  -0.1026   0.0053   1.0000
  12.000   1.7205   0.02682   0.02085  -0.0994   0.0052   1.0000
  12.250   1.7277   0.02793   0.02206  -0.0963   0.0051   1.0000
  12.500   1.7341   0.02914   0.02339  -0.0932   0.0049   1.0000
  12.750   1.7396   0.03047   0.02484  -0.0902   0.0048   1.0000
  13.000   1.7438   0.03196   0.02644  -0.0872   0.0047   1.0000
  13.250   1.7473   0.03356   0.02817  -0.0844   0.0046   1.0000
  13.500   1.7494   0.03533   0.03005  -0.0816   0.0045   1.0000
  13.750   1.7505   0.03728   0.03212  -0.0790   0.0044   1.0000
  14.000   1.7504   0.03940   0.03437  -0.0766   0.0043   1.0000
  14.250   1.7483   0.04181   0.03691  -0.0743   0.0042   1.0000
  14.500   1.7451   0.04447   0.03970  -0.0723   0.0041   1.0000
  14.750   1.7395   0.04751   0.04289  -0.0705   0.0041   1.0000
  15.000   1.7321   0.05094   0.04647  -0.0690   0.0040   1.0000
  15.250   1.7222   0.05490   0.05057  -0.0680   0.0040   1.0000
  15.500   1.7098   0.05940   0.05522  -0.0676   0.0039   1.0000
  15.750   1.6951   0.06455   0.06054  -0.0677   0.0039   1.0000
  16.000   1.6795   0.07026   0.06641  -0.0686   0.0038   1.0000
  16.250   1.6611   0.07680   0.07311  -0.0702   0.0038   1.0000
  16.500   1.6409   0.08402   0.08051  -0.0726   0.0038   1.0000
  16.750   1.6214   0.09160   0.08824  -0.0756   0.0038   1.0000
  17.000   1.6000   0.09990   0.09670  -0.0792   0.0038   1.0000
  17.250   1.5774   0.10873   0.10570  -0.0833   0.0038   1.0000
  17.500   1.5541   0.11808   0.11521  -0.0880   0.0038   1.0000
<< Back to ARA-D 6% AIRFOIL (arad6-il)

Polar data table (+)

Polar graphs


<< Back to ARA-D 6% AIRFOIL (arad6-il)