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=9


Details Polar file
Airfoil: NASA SC(2)-0706 AIRFOIL (sc20706-il)
Reynolds number: 200,000
Max Cl/Cd: 40.6 at α=1.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20706-il-200000.txt
Download as CSV file: xf-sc20706-il-200000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5869   0.08833   0.08473  -0.0161   1.0000   0.0423
  -9.000  -0.5937   0.08451   0.08097  -0.0174   1.0000   0.0433
  -8.750  -0.5994   0.07963   0.07617  -0.0216   1.0000   0.0443
  -8.500  -0.6000   0.07245   0.06893  -0.0320   1.0000   0.0451
  -8.250  -0.5961   0.06706   0.06317  -0.0405   1.0000   0.0464
  -8.000  -0.5914   0.06019   0.05603  -0.0447   1.0000   0.0471
  -7.750  -0.5833   0.05539   0.05147  -0.0438   1.0000   0.0491
  -7.500  -0.5693   0.05300   0.04908  -0.0438   1.0000   0.0530
  -7.250  -0.5428   0.04677   0.04204  -0.0513   1.0000   0.0608
  -7.000  -0.5255   0.04288   0.03829  -0.0516   1.0000   0.0630
  -6.750  -0.5019   0.03979   0.03503  -0.0531   1.0000   0.0676
  -6.500  -0.4712   0.03557   0.03033  -0.0568   1.0000   0.0766
  -6.250  -0.4236   0.02710   0.02075  -0.0600   1.0000   0.0429
  -6.000  -0.3835   0.02224   0.01494  -0.0620   1.0000   0.0355
  -5.750  -0.3540   0.02022   0.01279  -0.0631   1.0000   0.0388
  -5.500  -0.3232   0.01842   0.01073  -0.0636   1.0000   0.0399
  -5.250  -0.2924   0.01682   0.00894  -0.0640   1.0000   0.0412
  -5.000  -0.2616   0.01558   0.00756  -0.0645   1.0000   0.0443
  -4.750  -0.2287   0.01428   0.00627  -0.0659   1.0000   0.0513
  -4.500  -0.1949   0.01336   0.00532  -0.0674   1.0000   0.0617
  -4.250  -0.1587   0.01244   0.00447  -0.0697   1.0000   0.0873
  -4.000  -0.0997   0.00978   0.00400  -0.0785   1.0000   0.6058
  -3.750  -0.0736   0.00982   0.00411  -0.0777   1.0000   0.6632
  -3.500  -0.0492   0.00995   0.00427  -0.0765   1.0000   0.7030
  -3.250  -0.0272   0.01015   0.00452  -0.0746   1.0000   0.7348
  -3.000  -0.0035   0.01032   0.00467  -0.0733   1.0000   0.7574
  -2.750   0.0168   0.01049   0.00486  -0.0710   1.0000   0.7765
  -2.500   0.0386   0.01063   0.00501  -0.0693   1.0000   0.7933
  -2.250   0.0607   0.01074   0.00513  -0.0677   1.0000   0.8082
  -2.000   0.0830   0.01081   0.00522  -0.0661   1.0000   0.8214
  -1.750   0.1066   0.01084   0.00526  -0.0651   1.0000   0.8322
  -1.500   0.1327   0.01085   0.00528  -0.0647   1.0000   0.8420
  -1.250   0.1581   0.01085   0.00532  -0.0642   1.0000   0.8512
  -1.000   0.1824   0.01084   0.00535  -0.0634   1.0000   0.8597
  -0.750   0.2087   0.01085   0.00541  -0.0632   1.0000   0.8690
  -0.500   0.2334   0.01085   0.00549  -0.0626   1.0000   0.8781
  -0.250   0.2576   0.01084   0.00556  -0.0618   1.0000   0.8868
   0.000   0.2837   0.01086   0.00566  -0.0616   1.0000   0.8961
   0.250   0.3084   0.01086   0.00578  -0.0611   1.0000   0.9059
   0.500   0.3319   0.01084   0.00591  -0.0602   1.0000   0.9159
   0.750   0.3557   0.01083   0.00604  -0.0595   1.0000   0.9270
   1.000   0.4147   0.01039   0.00583  -0.0658   0.9766   0.9330
   1.250   0.4896   0.01206   0.00455  -0.0732   0.1946   0.9310
   1.500   0.5060   0.01340   0.00524  -0.0709   0.0743   0.9419
   1.750   0.5265   0.01396   0.00576  -0.0692   0.0583   0.9542
   2.000   0.5471   0.01473   0.00658  -0.0675   0.0511   0.9693
   2.250   0.5744   0.01574   0.00755  -0.0674   0.0440   1.0000
   2.500   0.6064   0.01724   0.00914  -0.0682   0.0399   1.0000
   2.750   0.6393   0.01893   0.01102  -0.0689   0.0380   1.0000
   3.000   0.6725   0.02113   0.01348  -0.0695   0.0378   1.0000
   3.250   0.7046   0.02406   0.01680  -0.0697   0.0395   1.0000
   3.500   0.7351   0.02651   0.01966  -0.0697   0.0382   1.0000
   4.000   0.7952   0.03748   0.03208  -0.0663   0.0717   1.0000
   4.250   0.8165   0.04103   0.03610  -0.0656   0.0635   1.0000
   4.500   0.8387   0.04403   0.03933  -0.0654   0.0590   1.0000
   4.750   0.8587   0.04813   0.04337  -0.0658   0.0567   1.0000
   5.000   0.8732   0.05262   0.04861  -0.0643   0.0513   1.0000
   5.250   0.8898   0.05653   0.05281  -0.0640   0.0476   1.0000
   5.500   0.9043   0.06008   0.05645  -0.0637   0.0453   1.0000
   5.750   0.9161   0.06538   0.06153  -0.0641   0.0432   1.0000
   6.000   0.9150   0.07274   0.06922  -0.0637   0.0423   1.0000
   6.250   0.9241   0.07647   0.07356  -0.0649   0.0391   1.0000
   6.500   0.9271   0.08193   0.07918  -0.0665   0.0374   1.0000
   6.750   0.9269   0.08732   0.08467  -0.0685   0.0363   1.0000
   7.000   0.9239   0.09272   0.09013  -0.0710   0.0355   1.0000
   7.250   0.9160   0.09790   0.09534  -0.0736   0.0351   1.0000
   7.500   0.9084   0.10380   0.10125  -0.0786   0.0348   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)