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)-0402 AIRFOIL (sc20402-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0402 AIRFOIL (sc20402-il)
Reynolds number: 200,000
Max Cl/Cd: 20.14 at α=0.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20402-il-200000.txt
Download as CSV file: xf-sc20402-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0402 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.6845   0.10555   0.10223   0.0276   1.0000   0.0166
  -8.250  -0.6806   0.10146   0.09815   0.0271   1.0000   0.0172
  -8.000  -0.6767   0.09761   0.09431   0.0255   1.0000   0.0177
  -7.750  -0.6721   0.09372   0.09046   0.0232   1.0000   0.0182
  -7.500  -0.6640   0.08945   0.08620   0.0194   1.0000   0.0187
  -7.250  -0.6534   0.08488   0.08164   0.0146   1.0000   0.0193
  -7.000  -0.6402   0.08008   0.07682   0.0092   1.0000   0.0200
  -6.750  -0.6246   0.07513   0.07183   0.0037   1.0000   0.0208
  -6.500  -0.6065   0.07011   0.06675  -0.0018   1.0000   0.0217
  -6.250  -0.5859   0.06508   0.06161  -0.0071   1.0000   0.0227
  -6.000  -0.5625   0.06011   0.05650  -0.0119   1.0000   0.0241
  -5.750  -0.5336   0.05546   0.05159  -0.0166   1.0000   0.0261
  -5.500  -0.4949   0.05314   0.04847  -0.0203   1.0000   0.0277
  -5.250  -0.4725   0.04661   0.04167  -0.0236   1.0000   0.0286
  -5.000  -0.4567   0.04129   0.03643  -0.0252   1.0000   0.0314
  -4.750  -0.4300   0.03779   0.03271  -0.0269   1.0000   0.0358
  -4.500  -0.3935   0.03546   0.02963  -0.0284   1.0000   0.0417
  -4.250  -0.3717   0.03087   0.02515  -0.0300   1.0000   0.0473
  -4.000  -0.3416   0.02786   0.02176  -0.0312   1.0000   0.0578
  -3.750  -0.3130   0.02522   0.01886  -0.0322   1.0000   0.0720
  -3.500  -0.2844   0.02287   0.01625  -0.0333   1.0000   0.0968
  -2.750  -0.1860   0.01697   0.00896  -0.0307   1.0000   0.0417
  -2.500  -0.1567   0.01485   0.00672  -0.0299   1.0000   0.0261
  -2.250  -0.1282   0.01303   0.00481  -0.0294   1.0000   0.0215
  -2.000  -0.0989   0.01152   0.00327  -0.0291   1.0000   0.0173
  -1.750  -0.0695   0.01057   0.00224  -0.0293   1.0000   0.0170
  -1.500  -0.0400   0.00989   0.00133  -0.0294   1.0000   0.0218
  -1.250  -0.0355   0.00642   0.00103  -0.0237   1.0000   1.0000
  -1.000  -0.0073   0.00642   0.00082  -0.0237   1.0000   1.0000
  -0.750   0.0205   0.00642   0.00073  -0.0238   1.0000   1.0000
  -0.500   0.0482   0.00643   0.00069  -0.0238   1.0000   1.0000
  -0.250   0.0758   0.00644   0.00072  -0.0239   1.0000   1.0000
   0.000   0.1032   0.00647   0.00080  -0.0239   1.0000   1.0000
   0.250   0.1307   0.00649   0.00096  -0.0238   1.0000   1.0000
   0.500   0.1734   0.01021   0.00170  -0.0269   0.0188   1.0000
   0.750   0.2014   0.01101   0.00270  -0.0266   0.0167   1.0000
   1.000   0.2291   0.01218   0.00390  -0.0260   0.0189   1.0000
   1.250   0.2568   0.01366   0.00543  -0.0253   0.0225   1.0000
   1.500   0.2847   0.01601   0.00782  -0.0248   0.0283   1.0000
   1.750   0.3180   0.01720   0.00945  -0.0229   0.0617   1.0000
   2.000   0.3604   0.01930   0.01252  -0.0200   0.1637   1.0000
   2.250   0.3859   0.02138   0.01466  -0.0201   0.1251   1.0000
   2.500   0.4103   0.02390   0.01714  -0.0203   0.0948   1.0000
   2.750   0.4363   0.02589   0.01944  -0.0204   0.0702   1.0000
   3.000   0.4622   0.02849   0.02229  -0.0205   0.0566   1.0000
   3.250   0.4895   0.03118   0.02541  -0.0204   0.0463   1.0000
   4.500   0.6008   0.05271   0.04802  -0.0235   0.0276   1.0000
   4.750   0.6376   0.05426   0.05044  -0.0246   0.0256   1.0000
   5.000   0.6639   0.05854   0.05498  -0.0268   0.0236   1.0000
   5.250   0.6863   0.06321   0.05984  -0.0295   0.0222   1.0000
   5.500   0.7065   0.06802   0.06479  -0.0326   0.0210   1.0000
   5.750   0.7247   0.07290   0.06978  -0.0361   0.0200   1.0000
   6.000   0.7407   0.07780   0.07476  -0.0398   0.0192   1.0000
   6.250   0.7545   0.08268   0.07971  -0.0436   0.0185   1.0000
   6.500   0.7660   0.08752   0.08458  -0.0473   0.0178   1.0000
   6.750   0.7753   0.09228   0.08937  -0.0506   0.0173   1.0000
   7.000   0.7821   0.09699   0.09409  -0.0531   0.0167   1.0000
   7.250   0.7854   0.10193   0.09903  -0.0534   0.0160   1.0000
   7.500   0.7854   0.10808   0.10518  -0.0544   0.0155   1.0000
   7.750   0.6724   0.09993   0.09709  -0.0511   0.0189   1.0000
   8.000   0.6717   0.10416   0.10130  -0.0523   0.0186   1.0000
<< Back to NASA SC(2)-0402 AIRFOIL (sc20402-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0402 AIRFOIL (sc20402-il)