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

FX 63-120 AIRFOIL (fx63120-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: FX 63-120 AIRFOIL (fx63120-il)
Reynolds number: 500,000
Max Cl/Cd: 125.74 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx63120-il-500000.txt
Download as CSV file: xf-fx63120-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 63-120 AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.2639   0.09582   0.09356  -0.0555   0.9932   0.0279
  -9.250  -0.2618   0.08548   0.08322  -0.0680   0.9880   0.0289
  -9.000  -0.2503   0.08053   0.07828  -0.0720   0.9855   0.0294
  -8.750  -0.2317   0.07801   0.07576  -0.0746   0.9818   0.0298
  -8.500  -0.2131   0.07531   0.07306  -0.0777   0.9767   0.0305
  -8.250  -0.1945   0.07096   0.06871  -0.0836   0.9730   0.0316
  -8.000  -0.2227   0.03431   0.03120  -0.1380   0.9383   0.0297
  -7.750  -0.1848   0.02750   0.02389  -0.1476   0.9346   0.0276
  -7.500  -0.1453   0.02211   0.01780  -0.1549   0.9285   0.0279
  -7.250  -0.1031   0.02008   0.01528  -0.1594   0.9236   0.0288
  -7.000  -0.0624   0.01740   0.01237  -0.1639   0.9192   0.0303
  -6.750  -0.0276   0.01654   0.01143  -0.1658   0.9110   0.0314
  -6.500   0.0103   0.01542   0.01011  -0.1683   0.9045   0.0325
  -6.250   0.0423   0.01456   0.00907  -0.1694   0.8952   0.0335
  -6.000   0.0768   0.01420   0.00850  -0.1708   0.8875   0.0346
  -5.750   0.1047   0.01284   0.00707  -0.1713   0.8781   0.0359
  -5.500   0.1359   0.01225   0.00641  -0.1721   0.8706   0.0370
  -5.250   0.1642   0.01181   0.00594  -0.1723   0.8617   0.0381
  -5.000   0.1949   0.01147   0.00551  -0.1729   0.8544   0.0397
  -4.750   0.2233   0.01113   0.00512  -0.1731   0.8459   0.0409
  -4.500   0.2538   0.01076   0.00464  -0.1736   0.8379   0.0419
  -4.250   0.2836   0.01020   0.00406  -0.1742   0.8282   0.0438
  -4.000   0.3137   0.00997   0.00377  -0.1747   0.8199   0.0460
  -3.750   0.3431   0.00976   0.00352  -0.1750   0.8117   0.0484
  -3.500   0.3739   0.00953   0.00320  -0.1756   0.8051   0.0511
  -3.250   0.4038   0.00929   0.00297  -0.1761   0.7974   0.0562
  -3.000   0.4345   0.00908   0.00273  -0.1766   0.7907   0.0676
  -2.750   0.4687   0.00831   0.00244  -0.1787   0.7836   0.2105
  -2.500   0.5009   0.00791   0.00243  -0.1800   0.7767   0.3636
  -2.250   0.5300   0.00792   0.00242  -0.1801   0.7700   0.3941
  -2.000   0.5588   0.00791   0.00241  -0.1802   0.7626   0.4145
  -1.750   0.5879   0.00794   0.00240  -0.1803   0.7563   0.4335
  -1.500   0.6165   0.00792   0.00241  -0.1803   0.7490   0.4499
  -1.250   0.6454   0.00795   0.00241  -0.1804   0.7423   0.4654
  -1.000   0.6741   0.00795   0.00244  -0.1805   0.7352   0.4818
  -0.750   0.7027   0.00797   0.00248  -0.1805   0.7282   0.5027
  -0.500   0.7313   0.00801   0.00252  -0.1805   0.7214   0.5188
  -0.250   0.7597   0.00803   0.00257  -0.1805   0.7139   0.5341
   0.000   0.7882   0.00810   0.00262  -0.1805   0.7071   0.5490
   0.250   0.8165   0.00811   0.00268  -0.1805   0.6992   0.5633
   0.500   0.8449   0.00820   0.00273  -0.1805   0.6920   0.5788
   0.750   0.8730   0.00822   0.00281  -0.1804   0.6837   0.5960
   1.000   0.9008   0.00831   0.00287  -0.1802   0.6752   0.6129
   1.250   0.9282   0.00835   0.00294  -0.1800   0.6639   0.6286
   1.500   0.9555   0.00842   0.00301  -0.1797   0.6527   0.6434
   1.750   0.9828   0.00853   0.00309  -0.1795   0.6430   0.6573
   2.000   1.0102   0.00861   0.00319  -0.1793   0.6326   0.6703
   2.250   1.0373   0.00870   0.00330  -0.1790   0.6224   0.6821
   2.750   1.0912   0.00892   0.00353  -0.1784   0.6002   0.7057
   3.000   1.1178   0.00904   0.00366  -0.1781   0.5881   0.7167
   3.250   1.1441   0.00918   0.00379  -0.1777   0.5748   0.7269
   3.500   1.1701   0.00935   0.00393  -0.1772   0.5593   0.7378
   3.750   1.1956   0.00952   0.00409  -0.1767   0.5421   0.7483
   4.000   1.2209   0.00971   0.00426  -0.1761   0.5243   0.7588
   4.250   1.2459   0.00994   0.00445  -0.1755   0.5053   0.7702
   4.500   1.2701   0.01021   0.00467  -0.1748   0.4857   0.7820
   4.750   1.2943   0.01048   0.00491  -0.1741   0.4680   0.7945
   5.000   1.3185   0.01074   0.00516  -0.1734   0.4512   0.8088
   5.250   1.3421   0.01100   0.00543  -0.1726   0.4337   0.8257
   5.750   1.3848   0.01153   0.00598  -0.1699   0.3954   0.8823
   6.000   1.4010   0.01170   0.00617  -0.1675   0.3767   1.0000
   6.250   1.4246   0.01214   0.00652  -0.1669   0.3558   1.0000
   6.500   1.4480   0.01257   0.00687  -0.1662   0.3319   1.0000
   6.750   1.4697   0.01311   0.00728  -0.1653   0.3065   1.0000
   7.000   1.4901   0.01373   0.00776  -0.1641   0.2788   1.0000
   7.250   1.5099   0.01439   0.00828  -0.1629   0.2516   1.0000
   7.500   1.5290   0.01507   0.00884  -0.1616   0.2279   1.0000
   7.750   1.5478   0.01575   0.00942  -0.1602   0.2066   1.0000
   8.000   1.5656   0.01647   0.01004  -0.1586   0.1867   1.0000
   8.250   1.5828   0.01718   0.01067  -0.1570   0.1652   1.0000
   8.500   1.5965   0.01804   0.01138  -0.1548   0.1427   1.0000
   8.750   1.6082   0.01891   0.01215  -0.1522   0.1218   1.0000
   9.000   1.6185   0.01985   0.01299  -0.1495   0.1056   1.0000
   9.250   1.6291   0.02079   0.01388  -0.1469   0.0935   1.0000
   9.500   1.6402   0.02171   0.01478  -0.1444   0.0844   1.0000
   9.750   1.6503   0.02272   0.01578  -0.1419   0.0765   1.0000
  10.000   1.6598   0.02380   0.01685  -0.1395   0.0696   1.0000
  10.250   1.6704   0.02483   0.01793  -0.1373   0.0640   1.0000
  10.500   1.6787   0.02607   0.01916  -0.1349   0.0588   1.0000
  10.750   1.6882   0.02726   0.02041  -0.1328   0.0549   1.0000
  11.000   1.6976   0.02851   0.02169  -0.1308   0.0515   1.0000
  11.250   1.7021   0.03020   0.02340  -0.1285   0.0481   1.0000
  11.500   1.7125   0.03148   0.02477  -0.1269   0.0460   1.0000
  11.750   1.7209   0.03296   0.02632  -0.1252   0.0438   1.0000
  12.000   1.7261   0.03479   0.02818  -0.1235   0.0416   1.0000
  12.250   1.7285   0.03696   0.03041  -0.1216   0.0396   1.0000
  12.500   1.7369   0.03862   0.03217  -0.1204   0.0380   1.0000
  12.750   1.7428   0.04057   0.03420  -0.1191   0.0363   1.0000
  13.000   1.7448   0.04300   0.03668  -0.1178   0.0346   1.0000
  13.250   1.7406   0.04621   0.03998  -0.1164   0.0331   1.0000
  13.500   1.7471   0.04832   0.04221  -0.1155   0.0319   1.0000
  13.750   1.7511   0.05078   0.04476  -0.1147   0.0305   1.0000
  14.000   1.7522   0.05364   0.04770  -0.1141   0.0292   1.0000
  14.250   1.7466   0.05749   0.05161  -0.1135   0.0279   1.0000
  14.500   1.7455   0.06089   0.05513  -0.1131   0.0268   1.0000
  14.750   1.7479   0.06391   0.05827  -0.1129   0.0256   1.0000
  15.000   1.7486   0.06724   0.06170  -0.1129   0.0246   1.0000
  15.250   1.7465   0.07107   0.06562  -0.1131   0.0236   1.0000
  15.500   1.7374   0.07605   0.07068  -0.1136   0.0227   1.0000
  15.750   1.7333   0.08043   0.07519  -0.1142   0.0220   1.0000
  16.000   1.7323   0.08441   0.07931  -0.1149   0.0212   1.0000
  16.250   1.7299   0.08871   0.08373  -0.1157   0.0205   1.0000
  16.500   1.7270   0.09317   0.08829  -0.1168   0.0198   1.0000
  16.750   1.7218   0.09808   0.09329  -0.1182   0.0192   1.0000
  17.000   1.7123   0.10379   0.09909  -0.1200   0.0187   1.0000
<< Back to FX 63-120 AIRFOIL (fx63120-il)

Polar data table (+)

Polar graphs


<< Back to FX 63-120 AIRFOIL (fx63120-il)