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NASA SC(2)-0706 AIRFOIL (sc20706-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0706 AIRFOIL (sc20706-il)
Reynolds number: 100,000
Max Cl/Cd: 33.26 at α=1.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20706-il-100000-n5.txt
Download as CSV file: xf-sc20706-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.6050   0.07650   0.07146  -0.0288   1.0000   0.0264
  -8.750  -0.6062   0.07127   0.06620  -0.0326   1.0000   0.0259
  -8.500  -0.6052   0.06594   0.06080  -0.0361   1.0000   0.0255
  -8.250  -0.6006   0.06054   0.05525  -0.0396   1.0000   0.0251
  -8.000  -0.5916   0.05503   0.04952  -0.0430   1.0000   0.0247
  -7.750  -0.5774   0.04950   0.04366  -0.0463   1.0000   0.0244
  -7.500  -0.5578   0.04405   0.03779  -0.0496   1.0000   0.0242
  -7.250  -0.5331   0.03894   0.03214  -0.0526   1.0000   0.0242
  -7.000  -0.5044   0.03457   0.02714  -0.0552   1.0000   0.0251
  -6.750  -0.4736   0.03162   0.02348  -0.0570   1.0000   0.0271
  -6.500  -0.4440   0.02784   0.01918  -0.0590   1.0000   0.0281
  -6.250  -0.4148   0.02529   0.01626  -0.0599   1.0000   0.0288
  -6.000  -0.3861   0.02328   0.01397  -0.0605   1.0000   0.0299
  -5.750  -0.3578   0.02161   0.01209  -0.0607   1.0000   0.0316
  -5.500  -0.3294   0.02017   0.01049  -0.0609   1.0000   0.0339
  -5.250  -0.3012   0.01919   0.00933  -0.0611   1.0000   0.0391
  -5.000  -0.2713   0.01793   0.00805  -0.0621   1.0000   0.0449
  -4.750  -0.2405   0.01699   0.00700  -0.0630   1.0000   0.0529
  -4.500  -0.2088   0.01613   0.00614  -0.0642   1.0000   0.0682
  -4.250  -0.1752   0.01518   0.00533  -0.0659   1.0000   0.1063
  -4.000  -0.1336   0.01317   0.00481  -0.0708   1.0000   0.4186
  -3.750  -0.1087   0.01283   0.00511  -0.0697   1.0000   0.6113
  -3.500  -0.0863   0.01293   0.00530  -0.0679   1.0000   0.6765
  -3.250  -0.0655   0.01309   0.00548  -0.0657   1.0000   0.7186
  -3.000  -0.0450   0.01323   0.00561  -0.0634   1.0000   0.7483
  -2.500  -0.0005   0.01336   0.00563  -0.0603   1.0000   0.7868
  -2.250   0.0237   0.01337   0.00560  -0.0593   1.0000   0.8002
  -2.000   0.0498   0.01336   0.00555  -0.0590   1.0000   0.8116
  -1.750   0.0739   0.01334   0.00553  -0.0582   1.0000   0.8212
  -1.500   0.0993   0.01332   0.00550  -0.0577   1.0000   0.8305
  -1.250   0.1256   0.01332   0.00551  -0.0574   1.0000   0.8404
  -1.000   0.1517   0.01332   0.00554  -0.0572   1.0000   0.8500
  -0.750   0.1769   0.01330   0.00558  -0.0566   1.0000   0.8587
  -0.500   0.2031   0.01332   0.00567  -0.0565   1.0000   0.8680
  -0.250   0.2289   0.01333   0.00578  -0.0561   1.0000   0.8782
   0.000   0.2530   0.01333   0.00588  -0.0555   1.0000   0.8882
   0.250   0.2778   0.01334   0.00603  -0.0550   1.0000   0.8987
   0.500   0.3027   0.01336   0.00620  -0.0545   1.0000   0.9099
   0.750   0.3266   0.01337   0.00642  -0.0539   1.0000   0.9226
   1.000   0.3495   0.01337   0.00662  -0.0530   1.0000   0.9374
   1.250   0.4154   0.01249   0.00606  -0.0597   0.9158   0.9387
   1.500   0.4920   0.01518   0.00552  -0.0677   0.1377   0.9357
   1.750   0.5139   0.01616   0.00610  -0.0666   0.0733   0.9487
   2.000   0.5385   0.01690   0.00691  -0.0658   0.0586   0.9660
   2.250   0.5666   0.01774   0.00778  -0.0660   0.0464   1.0000
   2.500   0.5954   0.01886   0.00901  -0.0663   0.0406   1.0000
   2.750   0.6233   0.02053   0.01067  -0.0666   0.0359   1.0000
   3.000   0.6538   0.02160   0.01196  -0.0670   0.0314   1.0000
   3.250   0.6839   0.02333   0.01386  -0.0673   0.0293   1.0000
   3.500   0.7140   0.02528   0.01605  -0.0676   0.0279   1.0000
   3.750   0.7431   0.02752   0.01859  -0.0676   0.0270   1.0000
   4.000   0.7709   0.03009   0.02155  -0.0675   0.0265   1.0000
   4.250   0.7968   0.03305   0.02504  -0.0670   0.0262   1.0000
   4.500   0.8195   0.03633   0.02873  -0.0665   0.0253   1.0000
   4.750   0.8396   0.04010   0.03310  -0.0654   0.0237   1.0000
   5.000   0.8595   0.04392   0.03751  -0.0641   0.0228   1.0000
   5.250   0.8757   0.04841   0.04248  -0.0629   0.0228   1.0000
   5.500   0.8892   0.05316   0.04766  -0.0619   0.0229   1.0000
   5.750   0.9001   0.05810   0.05296  -0.0610   0.0232   1.0000
   6.000   0.9082   0.06320   0.05835  -0.0604   0.0235   1.0000
   6.250   0.9132   0.06842   0.06380  -0.0599   0.0239   1.0000
   6.500   0.9172   0.07343   0.06895  -0.0596   0.0247   1.0000
   6.750   0.9156   0.08301   0.07908  -0.0624   0.0293   1.0000
   7.000   0.9113   0.08919   0.08537  -0.0645   0.0312   1.0000
   7.250   0.9054   0.09458   0.09080  -0.0661   0.0329   1.0000
   7.750   0.7570   0.09697   0.09336  -0.0584   0.0296   1.0000
   8.000   0.7472   0.10301   0.09937  -0.0611   0.0308   1.0000
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