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BOEING VERTOL V13006-.7 AIRFOIL (v13006-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: BOEING VERTOL V13006-.7 AIRFOIL (v13006-il)
Reynolds number: 200,000
Max Cl/Cd: 41.05 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-v13006-il-200000-n5.txt
Download as CSV file: xf-v13006-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING VERTOL V13006-.7 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.6344   0.09657   0.09302   0.0180   1.0000   0.0235
  -8.250  -0.6329   0.09234   0.08881   0.0151   1.0000   0.0239
  -8.000  -0.6320   0.08790   0.08441   0.0116   1.0000   0.0242
  -7.750  -0.6299   0.08295   0.07947   0.0063   1.0000   0.0245
  -7.500  -0.6236   0.07761   0.07410   0.0009   1.0000   0.0249
  -7.250  -0.6150   0.07226   0.06868  -0.0039   1.0000   0.0254
  -7.000  -0.5945   0.06714   0.06328  -0.0104   1.0000   0.0277
  -6.750  -0.5814   0.06277   0.05873  -0.0122   1.0000   0.0278
  -6.500  -0.5670   0.05868   0.05442  -0.0135   1.0000   0.0279
  -6.000  -0.5467   0.04340   0.03854  -0.0151   1.0000   0.0201
  -5.750  -0.5286   0.04050   0.03552  -0.0151   1.0000   0.0188
  -5.500  -0.5103   0.03670   0.03145  -0.0150   1.0000   0.0182
  -5.250  -0.4901   0.03306   0.02748  -0.0146   1.0000   0.0178
  -5.000  -0.4685   0.02966   0.02370  -0.0140   1.0000   0.0174
  -4.750  -0.4458   0.02656   0.02018  -0.0132   1.0000   0.0172
  -4.500  -0.4221   0.02386   0.01707  -0.0123   1.0000   0.0172
  -4.250  -0.3975   0.02159   0.01443  -0.0115   1.0000   0.0175
  -4.000  -0.3725   0.01973   0.01223  -0.0107   1.0000   0.0179
  -3.750  -0.3467   0.01868   0.01090  -0.0099   1.0000   0.0193
  -3.500  -0.3219   0.01710   0.00911  -0.0093   1.0000   0.0204
  -3.250  -0.2970   0.01587   0.00778  -0.0086   1.0000   0.0209
  -3.000  -0.2721   0.01487   0.00674  -0.0079   1.0000   0.0216
  -2.750  -0.2473   0.01404   0.00588  -0.0072   1.0000   0.0226
  -2.500  -0.2227   0.01330   0.00512  -0.0065   1.0000   0.0238
  -2.250  -0.1982   0.01270   0.00448  -0.0057   1.0000   0.0257
  -2.000  -0.1734   0.01228   0.00406  -0.0051   1.0000   0.0284
  -1.750  -0.1491   0.01157   0.00334  -0.0044   1.0000   0.0317
  -1.500  -0.1241   0.01118   0.00292  -0.0038   1.0000   0.0362
  -1.250  -0.0990   0.01081   0.00254  -0.0033   1.0000   0.0456
  -1.000  -0.0741   0.01029   0.00230  -0.0029   1.0000   0.1033
  -0.750  -0.0517   0.00768   0.00222  -0.0026   0.9936   0.7416
  -0.500   0.0069   0.00733   0.00231  -0.0076   0.9968   1.0000
  -0.250   0.0476   0.00733   0.00220  -0.0104   0.9788   1.0000
   0.000   0.0882   0.00732   0.00211  -0.0131   0.9514   1.0000
   0.250   0.1273   0.00731   0.00200  -0.0152   0.9093   1.0000
   0.500   0.1584   0.00733   0.00189  -0.0154   0.8599   1.0000
   0.750   0.1844   0.00742   0.00182  -0.0145   0.8120   1.0000
   1.000   0.2087   0.00756   0.00177  -0.0134   0.7631   1.0000
   1.250   0.2326   0.00776   0.00176  -0.0122   0.7106   1.0000
   1.500   0.2566   0.00801   0.00177  -0.0111   0.6564   1.0000
   1.750   0.2810   0.00828   0.00181  -0.0103   0.6037   1.0000
   2.000   0.3057   0.00855   0.00189  -0.0095   0.5538   1.0000
   2.250   0.3304   0.00887   0.00200  -0.0088   0.4991   1.0000
   2.500   0.3542   0.00937   0.00211  -0.0081   0.4163   1.0000
   2.750   0.3784   0.00988   0.00225  -0.0076   0.3364   1.0000
   3.000   0.4035   0.01029   0.00244  -0.0072   0.2867   1.0000
   3.250   0.4288   0.01069   0.00267  -0.0068   0.2431   1.0000
   3.500   0.4541   0.01113   0.00289  -0.0065   0.1900   1.0000
   3.750   0.4791   0.01167   0.00318  -0.0062   0.1340   1.0000
   4.000   0.5024   0.01273   0.00371  -0.0058   0.0416   1.0000
   4.250   0.5278   0.01333   0.00433  -0.0054   0.0298   1.0000
   4.500   0.5536   0.01386   0.00496  -0.0049   0.0261   1.0000
   4.750   0.5792   0.01440   0.00560  -0.0045   0.0242   1.0000
   5.000   0.6045   0.01500   0.00629  -0.0040   0.0229   1.0000
   5.250   0.6295   0.01564   0.00702  -0.0036   0.0213   1.0000
   5.500   0.6533   0.01656   0.00798  -0.0031   0.0198   1.0000
   5.750   0.6777   0.01736   0.00888  -0.0026   0.0191   1.0000
   6.000   0.7017   0.01832   0.00993  -0.0020   0.0186   1.0000
   6.250   0.7254   0.01940   0.01112  -0.0014   0.0182   1.0000
   6.500   0.7490   0.02062   0.01248  -0.0007   0.0178   1.0000
   6.750   0.7725   0.02199   0.01399  -0.0001   0.0174   1.0000
   7.000   0.7956   0.02352   0.01572   0.0006   0.0172   1.0000
   7.250   0.8182   0.02527   0.01771   0.0013   0.0170   1.0000
   7.500   0.8399   0.02722   0.01994   0.0021   0.0168   1.0000
   7.750   0.8611   0.02891   0.02190   0.0027   0.0162   1.0000
   8.000   0.8814   0.03048   0.02368   0.0032   0.0156   1.0000
   8.250   0.8998   0.03251   0.02591   0.0037   0.0150   1.0000
   8.500   0.9147   0.03552   0.02928   0.0045   0.0148   1.0000
   8.750   0.9259   0.03912   0.03329   0.0053   0.0147   1.0000
   9.000   0.9338   0.04293   0.03752   0.0060   0.0146   1.0000
   9.250   0.9379   0.04694   0.04195   0.0066   0.0146   1.0000
   9.500   0.9368   0.05131   0.04669   0.0068   0.0146   1.0000
   9.750   0.9302   0.05592   0.05161   0.0065   0.0146   1.0000
  10.250   0.9012   0.06582   0.06197   0.0019   0.0147   1.0000
  10.500   0.8313   0.09614   0.09260  -0.0241   0.0161   1.0000
  10.750   0.8165   0.10531   0.10172  -0.0290   0.0166   1.0000
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