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

NASA SC(2)-0706 AIRFOIL (sc20706-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0706 AIRFOIL (sc20706-il)
Reynolds number: 200,000
Max Cl/Cd: 44.13 at α=0.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20706-il-200000-n5.txt
Download as CSV file: xf-sc20706-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0706 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.5976   0.08878   0.08510  -0.0162   1.0000   0.0154
  -9.500  -0.6023   0.08414   0.08050  -0.0183   1.0000   0.0157
  -9.250  -0.6099   0.07969   0.07612  -0.0199   1.0000   0.0158
  -9.000  -0.6164   0.07429   0.07078  -0.0236   1.0000   0.0157
  -8.750  -0.6187   0.06705   0.06351  -0.0319   1.0000   0.0155
  -8.500  -0.6174   0.06040   0.05673  -0.0373   1.0000   0.0156
  -8.250  -0.6108   0.05405   0.05017  -0.0416   1.0000   0.0157
  -8.000  -0.5990   0.04779   0.04362  -0.0453   1.0000   0.0154
  -7.750  -0.5808   0.04149   0.03691  -0.0490   1.0000   0.0151
  -7.500  -0.5560   0.03540   0.03029  -0.0526   1.0000   0.0149
  -7.250  -0.5262   0.03005   0.02432  -0.0558   1.0000   0.0149
  -7.000  -0.4944   0.02593   0.01957  -0.0582   1.0000   0.0152
  -6.750  -0.4630   0.02297   0.01608  -0.0599   1.0000   0.0158
  -6.500  -0.4328   0.02088   0.01359  -0.0608   1.0000   0.0165
  -6.250  -0.4036   0.01943   0.01181  -0.0615   1.0000   0.0176
  -6.000  -0.3742   0.01783   0.01009  -0.0625   1.0000   0.0197
  -5.750  -0.3445   0.01663   0.00876  -0.0632   1.0000   0.0208
  -5.500  -0.3142   0.01558   0.00761  -0.0640   1.0000   0.0222
  -5.250  -0.2830   0.01468   0.00660  -0.0650   1.0000   0.0244
  -5.000  -0.2519   0.01402   0.00579  -0.0660   1.0000   0.0268
  -4.750  -0.2195   0.01330   0.00507  -0.0675   1.0000   0.0348
  -4.500  -0.1877   0.01276   0.00450  -0.0687   1.0000   0.0438
  -4.250  -0.1561   0.01231   0.00407  -0.0698   1.0000   0.0637
  -4.000  -0.1220   0.01169   0.00370  -0.0718   1.0000   0.1221
  -3.750  -0.0785   0.01042   0.00335  -0.0766   1.0000   0.3605
  -3.500  -0.0427   0.00982   0.00344  -0.0788   1.0000   0.5510
  -3.250  -0.0145   0.00974   0.00357  -0.0788   1.0000   0.6239
  -2.750   0.0389   0.00981   0.00374  -0.0780   1.0000   0.6926
  -2.500   0.0656   0.00987   0.00382  -0.0776   1.0000   0.7133
  -2.250   0.0920   0.00993   0.00390  -0.0772   1.0000   0.7293
  -2.000   0.1184   0.01000   0.00400  -0.0768   1.0000   0.7432
  -1.750   0.1448   0.01006   0.00409  -0.0765   1.0000   0.7540
  -1.500   0.1715   0.01012   0.00418  -0.0762   1.0000   0.7625
  -1.250   0.1989   0.01018   0.00427  -0.0762   1.0000   0.7712
  -1.000   0.2252   0.01025   0.00441  -0.0758   1.0000   0.7783
  -0.750   0.2520   0.01033   0.00456  -0.0757   1.0000   0.7861
  -0.500   0.2785   0.01042   0.00473  -0.0754   1.0000   0.7932
  -0.250   0.3054   0.01051   0.00492  -0.0753   1.0000   0.8002
   0.250   0.4054   0.00974   0.00443  -0.0841   0.9483   0.8119
   0.500   0.4770   0.01081   0.00361  -0.0918   0.4426   0.8146
   0.750   0.4954   0.01231   0.00404  -0.0904   0.1912   0.8211
   1.000   0.5188   0.01324   0.00448  -0.0898   0.0770   0.8273
   1.250   0.5440   0.01378   0.00495  -0.0891   0.0507   0.8337
   1.500   0.5697   0.01428   0.00544  -0.0886   0.0377   0.8404
   1.750   0.5949   0.01480   0.00605  -0.0879   0.0313   0.8470
   2.000   0.6200   0.01553   0.00679  -0.0874   0.0248   0.8533
   2.250   0.6444   0.01622   0.00764  -0.0864   0.0229   0.8593
   2.500   0.6697   0.01711   0.00864  -0.0857   0.0210   0.8661
   2.750   0.6934   0.01810   0.00975  -0.0846   0.0197   0.8722
   3.000   0.7191   0.01930   0.01108  -0.0839   0.0188   0.8791
   3.250   0.7431   0.02060   0.01256  -0.0827   0.0182   0.8852
   3.750   0.7902   0.02397   0.01637  -0.0809   0.0157   0.9000
   4.000   0.8143   0.02561   0.01838  -0.0798   0.0150   0.9078
   4.250   0.8363   0.02784   0.02106  -0.0782   0.0145   0.9156
   4.500   0.8564   0.03074   0.02448  -0.0762   0.0142   0.9240
   4.750   0.8740   0.03428   0.02856  -0.0739   0.0139   0.9341
   5.000   0.8883   0.03831   0.03311  -0.0713   0.0139   0.9474
   5.250   0.9003   0.04271   0.03800  -0.0687   0.0140   0.9912
   5.500   0.9146   0.04804   0.04378  -0.0674   0.0141   1.0000
   5.750   0.9254   0.05348   0.04958  -0.0662   0.0144   1.0000
   6.000   0.9340   0.05911   0.05553  -0.0654   0.0145   1.0000
   6.250   0.9417   0.06501   0.06172  -0.0649   0.0141   1.0000
   6.500   0.9463   0.07135   0.06831  -0.0652   0.0134   1.0000
   6.750   0.9472   0.07743   0.07457  -0.0662   0.0130   1.0000
   7.000   0.9443   0.08356   0.08083  -0.0679   0.0131   1.0000
   7.250   0.9378   0.08959   0.08695  -0.0702   0.0132   1.0000
   7.500   0.9272   0.09538   0.09278  -0.0731   0.0134   1.0000
<< Back to NASA SC(2)-0706 AIRFOIL (sc20706-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0706 AIRFOIL (sc20706-il)