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

D.G.A. 1182 (dga1182-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: D.G.A. 1182 (dga1182-il)
Reynolds number: 50,000
Max Cl/Cd: 22.65 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-dga1182-il-50000.txt
Download as CSV file: xf-dga1182-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: D.G.A. 1182                                     
  
 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  -0.6862   0.13502   0.12883   0.0300   1.0000   0.1868
 -10.000  -0.6714   0.12975   0.12357   0.0311   1.0000   0.1952
  -9.750  -0.6916   0.12800   0.12194   0.0268   1.0000   0.2006
  -9.500  -0.6771   0.12292   0.11686   0.0282   1.0000   0.2124
  -9.250  -0.6685   0.11854   0.11251   0.0287   1.0000   0.2239
  -9.000  -0.6651   0.11470   0.10872   0.0287   1.0000   0.2374
  -8.750  -0.6694   0.11157   0.10567   0.0281   1.0000   0.2537
  -8.500  -0.6560   0.10792   0.10204   0.0305   1.0000   0.2790
  -8.250  -0.6401   0.10504   0.09914   0.0337   1.0000   0.3139
  -8.000  -0.6415   0.10321   0.09737   0.0352   1.0000   0.3446
  -7.750  -0.6348   0.10030   0.09451   0.0371   1.0000   0.3727
  -7.500  -0.6267   0.09690   0.09116   0.0390   1.0000   0.3997
  -7.250  -0.6212   0.09356   0.08787   0.0407   1.0000   0.4267
  -7.000  -0.6190   0.09050   0.08488   0.0428   1.0000   0.4552
  -6.500  -0.5928   0.08268   0.07712   0.0467   1.0000   0.5128
  -5.500  -0.6154   0.05423   0.04839   0.0017   1.0000   0.3048
  -5.250  -0.5723   0.04418   0.03635  -0.0117   1.0000   0.1700
  -5.000  -0.5423   0.04080   0.03224  -0.0114   1.0000   0.1393
  -4.750  -0.5139   0.03771   0.02843  -0.0107   1.0000   0.1215
  -4.500  -0.4860   0.03486   0.02514  -0.0098   1.0000   0.1126
  -4.250  -0.4566   0.03243   0.02201  -0.0085   1.0000   0.1049
  -4.000  -0.4283   0.02994   0.01931  -0.0076   1.0000   0.1030
  -3.750  -0.3997   0.02777   0.01695  -0.0067   1.0000   0.1036
  -3.500  -0.3731   0.02576   0.01502  -0.0059   1.0000   0.1110
  -3.250  -0.3454   0.02416   0.01335  -0.0046   1.0000   0.1166
  -3.000  -0.3196   0.02267   0.01192  -0.0034   1.0000   0.1244
  -2.750  -0.2937   0.02145   0.01069  -0.0025   1.0000   0.1406
  -2.500  -0.2682   0.01988   0.00920  -0.0017   1.0000   0.1940
  -2.250  -0.2593   0.01714   0.00805   0.0028   1.0000   0.5549
  -2.000  -0.2533   0.01607   0.00800   0.0097   1.0000   0.7079
  -1.750  -0.2429   0.01615   0.00853   0.0169   1.0000   0.8229
  -1.500  -0.2160   0.01595   0.00817   0.0175   1.0000   0.8574
  -1.250  -0.1826   0.01570   0.00773   0.0163   1.0000   0.8865
  -1.000  -0.1389   0.01548   0.00731   0.0130   1.0000   0.9195
  -0.750  -0.0828   0.01530   0.00694   0.0068   1.0000   0.9592
  -0.500  -0.0261   0.01508   0.00653  -0.0002   1.0000   1.0000
  -0.250  -0.0109   0.01492   0.00623   0.0010   1.0000   1.0000
   0.000   0.0067   0.01487   0.00605   0.0022   1.0000   1.0000
   0.250   0.0254   0.01487   0.00597   0.0034   1.0000   1.0000
   0.500   0.0446   0.01492   0.00596   0.0045   1.0000   1.0000
   0.750   0.0641   0.01499   0.00601   0.0055   1.0000   1.0000
   1.000   0.0838   0.01510   0.00611   0.0065   1.0000   1.0000
   1.250   0.1034   0.01525   0.00629   0.0074   1.0000   1.0000
   1.500   0.1231   0.01545   0.00652   0.0083   1.0000   1.0000
   1.750   0.1721   0.01566   0.00685   0.0034   0.9783   1.0000
   2.000   0.2291   0.01578   0.00714  -0.0025   0.9428   1.0000
   2.250   0.2781   0.01576   0.00730  -0.0059   0.8844   1.0000
   2.500   0.3160   0.01578   0.00740  -0.0068   0.8264   1.0000
   2.750   0.3453   0.01589   0.00750  -0.0056   0.7556   1.0000
   3.000   0.3671   0.01621   0.00744  -0.0024   0.6318   1.0000
   3.250   0.3850   0.01707   0.00764   0.0010   0.4743   1.0000
   3.500   0.4023   0.01869   0.00819   0.0031   0.3433   1.0000
   3.750   0.4259   0.01983   0.00908   0.0041   0.2654   1.0000
   4.000   0.4502   0.02066   0.00953   0.0049   0.2245   1.0000
   4.250   0.4761   0.02227   0.01079   0.0056   0.2072   1.0000
   4.500   0.5046   0.02494   0.01311   0.0061   0.1901   1.0000
   4.750   0.5337   0.02803   0.01611   0.0064   0.1809   1.0000
   5.000   0.5618   0.02978   0.01835   0.0070   0.1789   1.0000
   5.250   0.5898   0.03191   0.02089   0.0076   0.1795   1.0000
   5.500   0.6184   0.03347   0.02319   0.0084   0.1844   1.0000
   5.750   0.6455   0.03610   0.02652   0.0089   0.1901   1.0000
   6.000   0.6699   0.03949   0.03044   0.0092   0.1918   1.0000
   6.250   0.6937   0.04592   0.03834   0.0070   0.2241   1.0000
   6.500   0.7061   0.05510   0.04844   0.0009   0.2585   1.0000
   6.750   0.7185   0.06253   0.05616  -0.0035   0.2798   1.0000
   7.250   0.6265   0.08605   0.07965  -0.0487   0.5190   1.0000
   7.500   0.6403   0.08891   0.08254  -0.0454   0.4737   1.0000
   7.750   0.6501   0.09212   0.08575  -0.0435   0.4431   1.0000
   8.000   0.6602   0.09546   0.08910  -0.0417   0.4139   1.0000
   8.250   0.6688   0.09894   0.09258  -0.0401   0.3876   1.0000
   8.500   0.6795   0.10267   0.09631  -0.0383   0.3612   1.0000
   8.750   0.6995   0.10637   0.10007  -0.0344   0.3212   1.0000
   9.000   0.7076   0.10975   0.10349  -0.0324   0.2939   1.0000
   9.250   0.6957   0.11194   0.10559  -0.0331   0.2734   1.0000
   9.500   0.7096   0.11456   0.10824  -0.0298   0.2381   1.0000
  10.000   0.7369   0.12129   0.11501  -0.0232   0.1808   1.0000
  10.250   0.7187   0.12477   0.11837  -0.0279   0.1774   1.0000
  10.500   0.7280   0.12970   0.12330  -0.0272   0.1683   1.0000
<< Back to D.G.A. 1182 (dga1182-il)

Polar data table (+)

Polar graphs


<< Back to D.G.A. 1182 (dga1182-il)