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

SC(2)-0714 Supercritical airfoil (coordinates from Raymer w/ one correction) (nasasc2-0714-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: SC(2)-0714 Supercritical airfoil (coordinates from Raymer w/ one correction) (nasasc2-0714-il)
Reynolds number: 50,000
Max Cl/Cd: 24.55 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-nasasc2-0714-il-50000.txt
Download as CSV file: xf-nasasc2-0714-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: SC(2)-0714 Supercritical airfoil (coordinates fr
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.5741   0.09717   0.08906  -0.0376   1.0000   0.2053
 -10.000  -0.5599   0.09380   0.08569  -0.0363   1.0000   0.2026
  -9.750  -0.6432   0.08046   0.07248  -0.0416   1.0000   0.1832
  -9.500  -0.7679   0.06836   0.06001  -0.0461   1.0000   0.1709
  -9.250  -0.7707   0.06373   0.05523  -0.0461   1.0000   0.1695
  -9.000  -0.7723   0.05900   0.05021  -0.0468   1.0000   0.1685
  -8.750  -0.7686   0.05445   0.04524  -0.0482   1.0000   0.1689
  -8.500  -0.7567   0.05035   0.04060  -0.0500   1.0000   0.1707
  -8.250  -0.7389   0.04675   0.03665  -0.0512   1.0000   0.1732
  -8.000  -0.7189   0.04415   0.03397  -0.0510   1.0000   0.1761
  -7.750  -0.6971   0.04183   0.03146  -0.0512   1.0000   0.1807
  -7.500  -0.6718   0.03944   0.02862  -0.0525   1.0000   0.1869
  -7.250  -0.6484   0.03732   0.02641  -0.0523   1.0000   0.1926
  -7.000  -0.6248   0.03570   0.02470  -0.0519   1.0000   0.1998
  -6.750  -0.6005   0.03409   0.02296  -0.0516   1.0000   0.2094
  -6.500  -0.5770   0.03280   0.02161  -0.0508   1.0000   0.2208
  -6.250  -0.5566   0.03165   0.02063  -0.0488   1.0000   0.2332
  -6.000  -0.5361   0.03064   0.01974  -0.0468   1.0000   0.2507
  -5.750  -0.5151   0.02961   0.01884  -0.0451   1.0000   0.2746
  -5.500  -0.4945   0.02843   0.01810  -0.0435   1.0000   0.3114
  -5.250  -0.4746   0.02693   0.01773  -0.0419   1.0000   0.3908
  -5.000  -0.4818   0.02874   0.02075  -0.0296   1.0000   0.4898
  -4.750  -0.4778   0.03276   0.02480  -0.0188   1.0000   0.5843
  -4.500  -0.4698   0.03585   0.02773  -0.0103   1.0000   0.6331
  -4.250  -0.4651   0.03813   0.02994  -0.0008   1.0000   0.6606
  -4.000  -0.4562   0.03963   0.03133   0.0064   1.0000   0.6908
  -3.750  -0.4454   0.04049   0.03206   0.0121   1.0000   0.7215
  -3.500  -0.4346   0.04092   0.03240   0.0185   1.0000   0.7476
  -3.250  -0.4227   0.04085   0.03223   0.0232   1.0000   0.7757
  -3.000  -0.4111   0.04052   0.03179   0.0273   1.0000   0.8042
  -2.750  -0.3975   0.03992   0.03109   0.0315   1.0000   0.8346
  -2.500  -0.3649   0.03894   0.02999   0.0335   1.0000   0.8744
  -2.250  -0.2490   0.03775   0.02847   0.0199   1.0000   0.9305
  -2.000  -0.1773   0.03664   0.02717   0.0108   1.0000   0.9546
  -1.750  -0.1359   0.03584   0.02628   0.0065   1.0000   0.9677
  -1.500  -0.0984   0.03519   0.02557   0.0027   1.0000   0.9773
  -1.250  -0.0667   0.03470   0.02504   0.0000   1.0000   0.9848
  -1.000  -0.0367   0.03433   0.02465  -0.0024   1.0000   0.9913
  -0.750  -0.0099   0.03405   0.02437  -0.0042   1.0000   0.9964
  -0.500   0.0094   0.03388   0.02422  -0.0047   1.0000   1.0000
  -0.250   0.0136   0.03372   0.02409  -0.0023   1.0000   1.0000
   0.000   0.0176   0.03359   0.02400   0.0000   1.0000   1.0000
   0.250   0.0212   0.03349   0.02394   0.0024   1.0000   1.0000
   0.500   0.0246   0.03341   0.02391   0.0048   1.0000   1.0000
   0.750   0.0275   0.03336   0.02392   0.0072   1.0000   1.0000
   1.000   0.0300   0.03334   0.02395   0.0096   1.0000   1.0000
   1.250   0.0322   0.03334   0.02402   0.0120   1.0000   1.0000
   1.500   0.0339   0.03337   0.02412   0.0144   1.0000   1.0000
   1.750   0.0353   0.03343   0.02425   0.0167   1.0000   1.0000
   2.000   0.0365   0.03352   0.02442   0.0188   1.0000   1.0000
   2.250   0.0377   0.03366   0.02465   0.0209   1.0000   1.0000
   2.500   0.0391   0.03387   0.02495   0.0227   1.0000   1.0000
   2.750   0.0649   0.03485   0.02605   0.0199   0.9931   1.0000
   3.000   0.1668   0.03705   0.02845   0.0048   0.9513   1.0000
   3.250   0.2677   0.03751   0.02917  -0.0078   0.9037   1.0000
   3.500   0.3751   0.03615   0.02815  -0.0188   0.8571   1.0000
   3.750   0.4503   0.03382   0.02617  -0.0234   0.8121   1.0000
   4.000   0.5433   0.02868   0.02149  -0.0263   0.7450   1.0000
   4.250   0.6652   0.02710   0.01704  -0.0315   0.3420   1.0000
   4.500   0.7021   0.02875   0.01818  -0.0332   0.2853   1.0000
   4.750   0.7431   0.03027   0.01939  -0.0355   0.2528   1.0000
   5.000   0.7748   0.03164   0.02070  -0.0363   0.2337   1.0000
   5.250   0.8029   0.03310   0.02202  -0.0366   0.2195   1.0000
   5.500   0.8211   0.03435   0.02353  -0.0351   0.2098   1.0000
   5.750   0.8424   0.03596   0.02511  -0.0344   0.2011   1.0000
   6.000   0.8550   0.03732   0.02677  -0.0320   0.1943   1.0000
   6.250   0.8733   0.03899   0.02847  -0.0307   0.1885   1.0000
   6.500   0.8855   0.04090   0.03059  -0.0286   0.1838   1.0000
   6.750   0.8938   0.04269   0.03281  -0.0258   0.1794   1.0000
   7.000   0.9079   0.04489   0.03531  -0.0242   0.1756   1.0000
   7.250   0.9267   0.04738   0.03796  -0.0237   0.1722   1.0000
   7.500   0.9459   0.05054   0.04123  -0.0235   0.1687   1.0000
   7.750   0.9504   0.05354   0.04479  -0.0214   0.1669   1.0000
   8.000   0.9542   0.05711   0.04885  -0.0196   0.1662   1.0000
   8.250   0.9570   0.06120   0.05334  -0.0183   0.1668   1.0000
   8.500   0.9579   0.06560   0.05809  -0.0172   0.1675   1.0000
   8.750   0.9572   0.07027   0.06305  -0.0164   0.1682   1.0000
   9.000   0.9580   0.07527   0.06827  -0.0162   0.1690   1.0000
   9.250   0.8533   0.08438   0.07811  -0.0133   0.1829   1.0000
   9.500   0.8654   0.09011   0.08389  -0.0150   0.1855   1.0000
<< Back to SC(2)-0714 Supercritical airfoil (coordinates from Raymer w/ one correction) (nasasc2-0714-il)

Polar data table (+)

Polar graphs


<< Back to SC(2)-0714 Supercritical airfoil (coordinates from Raymer w/ one correction) (nasasc2-0714-il)